Method for forming a wear-resistant hard-face contact area on a workpiece, such as a gas turbine engine part

ABSTRACT

A method for forming a wear-resistant hardfaced contact area on the shroud section of a gas turbine engine blade. A predetermined contact area of a shroud section of a gas turbine engine blade is selectively coated with a high-density hardface coating material. The hardface coating material is capable of forming a diffusion boundary between the hardface coating material and the shroud section. A hot isostatic heat treatment process is performed to form the diffusion boundary between the hardface coating material and the shroud section to form a wear-resistant hardfaced contact area diffusion bonded to the shroud section. Depending on the coating process, and the necessity for doing so, the predetermined contact area can be masked off before the step of selectively coating. A sintering heat treatment can be perfomed before the step of performing the hot isostatic heat treatment to limit the occurrence bubbles on the surface of the hardface coating material after the isostatic heat treatment step. The sintering heat treatment may be performed at a temperature substantially the same as the temperature of the hot isostatic heat treatment. The hardface coating material may comprise an alloy with substantially no oxide forming constituents so as to avoid the formation of oxide inclusions in the coating material.

BACKGROUND OF THE INVENTION

The present invention pertains to methods for salvaging cast articles,such as turbine engine airfoil parts. More particularly, the presentinvention pertains to a method for correcting the dimensionalcharacteristics of a cast article; a method for applying a protectivecoating to a metal article; and a method for repairing a turbine engineairfoil part.

Airfoil parts, such as blades and vanes, are critical components in thegas turbine engines that are used to power jet aircraft or for thegeneration of electricity. Each airfoil part is an individual unithaving a root or attachment section and an airfoil section. The airfoilsection has specific cordal and length dimensions that define theairfoil characteristics of the part. The root section is engaged withand held by a housing member. A plurality of the airfoil parts are thusassembled with the housing member to form a disc or ring. Blades, whichduring operation are rotating part, are assembled into and disc. Vane,which remain stationary, are assembled into a nozzle or vane ring. Inthe operating gas turbine engine the assembled rings and discs,determine the path of the intake, combustion and exhaust gases that flowthrough the engine.

The airfoil part may be either a rotating component or a non-rotatingcomponent of the gas turbine engine. If the part is a rotatingcomponent, during operation of the turbine engine the part is subjectedto centrifugal forces that exert deforming stresses. These deformingstresses cause creep rupture and fatigue problems that can result in thefailure of the part. Non-rotating components, such as vanes, are notsubjected to centrifugal forces that exert deforming stresses. However,like the rotating parts, these parts are subjected to other deformationsuch as from hot gas erosion and/or foreign particle strikes. Thisdeformation results in the alteration of the dimensions of the airfoilsection. The alteration of the dimensions of the airfoil section candetrimentally modify the airflow through the gas turbine engine which iscritical to the engine's performance.

An example of a non-rotating airfoil part is the 2nd stage vane of thePratt & Whitney JT8D model 1 through 17R gas turbine engine. This partis manufactured by the “lost wax” or “investment casting” process. Thevane is cast from one of several highly alloyed nickel or cobalt-basematerials. As a new part in a new gas turbine engine, or as a new sparepart in an overhauled engine, it begins its life cycle with a protectivediffusion coating on its airfoil surfaces and a wear coating on surfacesknown to have excessive wear patterns.

When the gas turbine engine is operating, the vane will see temperaturesof about 1500 degree F. Since the vane does not rotate and thus is notsubject to creep rupture, its demise is most often influenced by thenumber of times it is repaired. The reason for this is the repairprocess itself.

The repair process consists of the following operations:

-   -   1). degrease, wash to remove engine carbon, etc.    -   2.) grit blast to remove wear coatings, and any sulfidation        which is present    -   3.) chemically remove the diffusion coating    -   4.) blend to remove nicks, dents, etc.    -   5.) weld, grind, polish etc.

The repair operations that remove metal by chemical stripping, gritblasting, blending and polishing shorten the life cycle of the vane. Thecoating removal is a major contributor because it is diffused into theparent metal. When certain minimum airfoil dimensions cannot be met thepart is deemed non-repairable and must be retired from service. Thus,there is a need for a method for repairing gas turbine engine airfoilparts that effectively and efficiently restores the airfoil dimensionsof the part.

On another front, during the manufacture of metal components a coatingoperation is performed to provide a coating material layer on thesurface of a component substrate. The coating material layer is formedto build-up the metal component to desired finished dimensions and toprovide the finished product with various surface attributes. Forexample, an oxide layer may be formed to provide a smooth, corrosionresistant surface. Also, a wear resistant coating, such as Carbide,Cobalt, or TiN is often formed on cutting tools to provide wearresistance.

Chemical Vapor Deposition is typically used to deposit a thin film wearresistant coating on a cutting tool substrate. For example, to increasethe service life of a drill bit, chemical vapor deposition can be usedto form a wear resistant coating of Cobalt on a high speed steel (HSS)cutting tool substrate. The bond between the substrate and coatingoccurs primarily through mechanical adhesion within a narrow bondinginterface. During use, the coating at the cutting surface of the cuttingtool is subjected to shearing forces resulting in flaking off thecoating of the tool substrate. The failure is likely to occur at thenarrow bonding interface.

FIG. 12(a) is a side view of a prior art tool bit coated with a wearresistant coating. In this case, the wear resistant coating may beapplied by the Chemical Vapor Deposition method so that the entire toolbit substrate receives an even thin film of a relatively hard material,such as Carbide, Cobalt or TiN. Since the coating adheres to the toolbit substrate mostly via a mechanical bond located at a boundaryinterface, flaking and chipping off the coating of the substrate islikely to occur during use, limiting the service life of the tool bit.FIG. 12(b) is a side view of a prior art tool bit having a fixed wearresistant cutting tip. In this case, a relatively hard metal cutting tipis fixed to the relatively soft tool bit substrate. The metal cuttingtip, which is typically comprised of a Carbide or Cobalt alloy, is fixedto the tool bit substrate by brazing. During extended use the tool bitis likely to fail at the relatively brittle brazed interface between themetal cutting tip and the tool substrate, and again, the useful servicelife of the tool bit is limited.

Another coating method, known as Conventional Plasma Spray uses a superheated inert gas to generate a plasma. Powder feedstock is introducedand carried to the workpiece by the plasma stream. Conventional plasmaspray coating methods deposit the coating material at relatively lowvelocity, resulting in voids being formed within the coating and in acoating density typically having a porosity of about 5.0%. Again, thebond between the substrate and the coating occurs primarily throughmechanical adhesion at a bonding interface, and if the coating issubjected to sufficient shearing forces it will flake off of theworkpiece substrate.

Another coating method, known as the Hyper Velocity Oxyfuel (HVOF)plasma thermal spray process is used to produce coatings that are nearlyabsent of voids. In fact, coatings can be produced nearly 100% dense,with a porosity of less than 0.5%. In HVOF thermal spraying, a fuel gasand oxygen are used to create a combustion flame at 2500 to 3 100° C.The combustion takes place at a very high chamber pressure and asupersonic gas stream forces the coating material through asmall-diameter barrel at very high particle velocities. The HVOF processresults in extremely dense, well-bonded coatings. Typically, HVOFcoatings can be formed nearly 100% dense, with a porosity of <0.5%. Thehigh particle velocities obtained using the HVOF process results inrelatively better bonding between the coating material and thesubstrate, as compared with other coating methods such as theConventional Plasma spray method or the Chemical Vapor Depositionmethod. However, the HVOF process also forms a bond between the coatingmaterial and the substrate that occurs primarily through mechanicaladhesion at a bonding interface.

Detonation Gun coating is another method that produces a relativelydense coating. Suspended powder is fed into a long tube along withoxygen and fuel gas. The mixture is ignited in a controlled explosion.High temperature and pressure is thus created to blast particles out ofthe end of the tube and toward the substrate to be coated.

An example of using HVOF or Detonation Gun coating techniques isdisclosed in U.S. Pat. No. 5,584,663, issued to Schell. This referencediscloses that the tips of turbine blades can be formed by melting andfusing a powder alloy. Preferrably, the blade tip is generated bydepositing molten metal alloy powder in multiple passes. Squealers atthe perimeter of the blade tip may be formed using methods such asDetonation Gun or HVOF spray methods. The forming step may be used togenerate a near-net shaped blade tip, and a subsequent machining stepmay be employed to generate the final or preferred shape of the bladetip.

Casting is a known method for forming metal components. Typically, asubstrate blank is cast to near-finished dimensions. Various machiningoperations, such as cutting, sanding and polishing are performed on thecast substrate blank to eventually obtain the metal component at desiredfinished dimensions. A cast metal component will typically have a numberof imperfections caused by voids and contaminants in the cast surfacestructure. The imperfections may be removed by machining away thesurface layer of the component, and/or by applying a surface coating.

The manufacture of metal components often entails costly operations toproduce products with the desired surface texture, material propertiesand dimensional tolerances. For example, a known process formanufacturing a metal component requires, among other steps, making acasting of the metal component, treating the metal component using a HotIsostatic Pressing (HIP) treatment process, and then machining the metalcomponent to remove surface imperfections and obtain the desireddimensional tolerances.

HIP treatment is used in the densification of cast metal components andas a diffusion bonding technique for consolidating powder metals. In theHIP treatment process, a part to be treated is raised to a hightemperature and isostatic pressure. Typically, the part is heated to0.6-0.8 times the melting point of the material comprising the part, andsubjected to pressures on the order of 0.2 to 0.5 times the yieldstrength of the material. Pressurization is achieved by pumping an inertgas, such as Argon, into a pressure vessel. Within the pressure vesselis a high temperature furnace, which heats the gas to the desiredtemperature. The temperature and pressure are held for a set length oftime, and then the gas is cooled and vented.

The HIP treatment process is used to produce near-net shaped components,reducing or eliminating the need for subsequent machining operations.Further, by precise control of the temperature, pressure and time of aHIP treatment schedule a particular microstructure for the treated partcan be obtained.

All casting processes must deal with problems that the wrought processesdo not encounter. Major among those are porosity and shrinkage that areminimized by elaborate gating techniques and other methods that increasecost and sometimes lower yield. However, the ability to produce anear-net or net shape is the motivating factor. In some cases, it ismore cost effective to intentionally cast the part not using elaborateand costly gating techniques and HIP treat the part to eliminate thesub-surface porosity. The surface of the part is then machined until thedense substrate is reached.

U.S. Pat. No. 5,156,321, issued to Liburdi et al and U.S. Pat. No.5,071,054, issued to Dzugan et al. are examples of methods that employthe HIP treatment process. Liburdi et al. discloses a technique torepair or join sections of a superalloy article. A powder matching thesuperalloy composition is sintered in its solid state to form a porousstructure in an area to be repaired or joined. A layer of matchingpowder, modified to incorporate melting point depressants, is added tothe surface of the sintered region. Liburdi discloses that the joint israised to a temperature where the modified layer melts while thesintered layer and base metal remain solid. The modified material flowsinto the sintered layer by capillary action resulting in a dense jointwith properties approaching those of the base metal. This referencediscloses that HIPing can be used as part of the heat treatment to closeany minor interior defects. Dzugan et al. discloses fabricating asuperalloy article by casting, and then refurbishing primary defects inthe surface of the cast piece. The defects are removed by grinding. Theaffected portions of the surface are first filled with a material thatis the same composition as the cast article. Then, a cladding powder isapplied to the surface through the use of a binder coat to obtain asmooth surface. The article is then heated to melt the cladding powder,and then cooled to solidify. Finally, the article is HIPed to achievefinal closure of the surface defects.

Metal alloy components, such as gas turbine parts such as blades andvanes, are often damaged during use. During operation, gas turbine partsare subjected to considerable degradation from high pressure andcentrifugal force in a hot corrosive atmosphere. The gas turbine partsalso sustain considerable damage due to impacts from foreign particles.This degradation results in a limited service life for these parts.Since they are costly to produce, various repair methods are employed torefurbish damaged gas turbine blades and vanes.

Some examples of methods employed to repair gas turbine blades and vanesinclude U.S. Pat. No. 4,291,448, issued to Cretella et al.; U.S. Pat.No. 4,028,787, issued to Cretella et al.; U.S. Pat. No. 4,866,828,issued to Fraser; and U.S. Pat. No. 4,837,389, issued to Shankar et al.

Cretella '448 discloses a process to restore turbine blade shrouds thathave lost their original dimensions due to wear while in service. Thisreference discloses using the known process of TIG welding worn portionsof a part with a weld wire of similar chemistry as the part substrate,followed by finish grinding. The part is then plasma sprayed with amaterial of similar chemistry to a net shape requiring little or nofinishing. The part is then sintered in an argon atmosphere. The plasmaspray process used in accordance with Cretella '448 results in a coatingporosity of about 5.0%. Even after sintering the coating remainsattached to the substrate and weld material only by mechanical bond atan interface bonding layer making the finished piece prone to chippingand flaking.

Cretella '787 discloses a process for restoring turbine vanes that havelost their original dimensions due to wear while in service. Again, aconventional plasma spray process is used to build up worn areas of thevane before performing a sintering operation in a vacuum or hydrogenfurnace. The porosity of the coating, and the interface bonding layer,results in a structure that is prone to chipping and flaking.

Fraser discloses a process to repair steam turbine blades or vanes thatutilize some method of connecting them together (i.e. lacing wire). Inaccordance with the method disclosed by Fraser, the area of a part thathas been distressed is removed and a new piece of like metal is weldedto the part. The lacing holes of the part are plug welded. The part isthen subjected to hot striking to return it to its original contour, andthe lacing holes are re-drilled.

Shankar et al. disclose a process for repairing gas turbine blades thatare distressed due to engine operation. A low-pressure plasma spraycoating is applied to the vanes and the part is re-contoured bygrinding. A coating of aluminum is then applied using a diffusioncoating process. Again, the conventional low-pressure plasma sprayprocess forms a mechanical bond at an interface boundary between thecoating and the substrate, resulting in a structure that is prone tofailure due to chipping and flaking.

Other examples of methods for repairing or improving the characteristicsof turbine engine airfoil parts include U.S. Pat. No. 5,451,142 issuedto Cetel et al.; U.S. Pat. No. 4,921,405, issued to Wilson; U.S. Pat.No. 4,145,481 issued to Gupta et al.; and U.S. Pat. No. 5,732,467 issuedto White et al.

Cetel discloses a turbine engine blade having a blade root with asurface having a thin zone of fine grains. A plasma spray technique isused to form a thin layer of material on the root or fir tree portion ofthe blade. The blade is then HIPed. After the HIP process, the blade issolution heat treated and then machined. This reference is directed to aprocess for modifying the root section of a turbine blade to improve themechanical properties of this area of the part. The root section isserrated and is attached to the disc by inserting the root serrationsinto matching serrations of the disc. The blade is normally produced, asrelating to chemistry and microstructure, to maximize the creep ruptureand high cycle fatigue properties of the airfoil which is exposed to thehot gas path. The root section of the part thus has those sameproperties as the airfoil section. However, the root section of theblade is exposed to stress of a type different than the airfoil section,usually referred to as low cycle fatigue. The root section experiencescolder operating temperatures than the airfoil section and is notdirectly in the path of the hot gases that flow through the engine.Also, the root section is subjected to metal to metal stress duringrotation resulting in low cycle fatigue cracking. Cetal is concernedwith treating only the fir tree or root portion of the blade to improveits mechanical properties. The root portion or a new or refurbishedblade is treated with a plasma spray process, HIPing, and a heattreatment and then machined. The blade is machined to remove materialfrom a high stress portion of the blade root. The material removed bythe machining operation is replaced by a zone of fine grains by a plasmaspray technique. The part is processed through a HIP cycle to densifythe deposit, and then a heat treatment cycle to enhance its properties.Finally, the root is machined back to the desired blueprint dimensionsand the part returned to service.

Wilson discloses a turbine engine blade having a single crystal bodyhaving an airfoil section and an attachment or root section. A layer ofpolycrystalline superalloy is applied to the attachment section,preferrably by plasma spraying. The coated blade is HIPed and thensolution heat-treated to optimize the polycrystalline microstructure.

Grupta discloses a process for producing high temperature corrosionresistant metal articles. A ductile metallic overlay is formed on thesurface of an article substrate, and an outer layer is applied over theoverlay. The article is then subjected to a HIP treatment to eliminateporosity and create an inter-diffusion between the outer layer theoverlay and the substrate.

U.S. Pat. No. 5,318,217, issued to Stinson et al, teaches a method forenhancing the structural integrity of a bond joint in a spray castarticle. In this reference, a molten metal is spray cast onto a metalsubstrate and treated by vacuum cleaning, boronizing and/or knurling toenhance the structural integrity of a diffusion bond joint formed by aHIP treatment. U.S. Pat. No. 5,211,776, issued to Weiman teaches that ametal and ceramic matrix composite can be formed by the successivebuilding up of layers. An optional HIP or diffusion annealing processcan be performed to improve the properties of the composite.

None of these prior attempts provide for the effective and efficientrestoration of the critical airfoil dimensions of a gas turbine engineairfoil part. Typically, an airfoil part will have to be discarded afterit has gone through a certain number of repair cycles. The stripping ofthe protective coating on the part during the repair process is a majorcontributing factor resulting in the discarding of the part. After anumber of repair cycles the part simply does not have the minimumdimensional characteristics necessary for it to perform its intendedfunction. Therefore, there is a need for a method for repairing gasturbine engine airfoil parts that effectively and efficiently restoresthe critical airfoil dimensions of the part.

Turbine engine airfoil parts, such as vanes, are manufactured to precisetolerances that determine the airflow characteristics for the part. Theclass of a turbine vane is the angular relationship between the airfoilsection and the inner and outer buttresses of the vane. This angularrelationship has a direct bearing on the angle of attack of the airfoilsection during the operation of the gas turbine engine. Over time, theangular relationship between the airfoil section and the inner and outerbuttresses of the vane may become altered due to, for example,deformation of the airfoil section from engine operation and repairprocesses and the like. Or, the particular angular relationship of theairfoil section and the inner and outer buttresses as originallymanufactured may need to be changed to improve engine performance. Inany event, there is a need for a method of restoring or reclassifying agas turbine engine airfoil part.

Airfoil parts, such as blades, are critical components in the gasturbine engines that are used to power jet aircraft or for thegeneration of electricity. As shown in FIG. 23(a), each blade 38 is anindividual unit having a shroud section 40 and an airfoil section 42.The airfoil section 42 has specific cordal and length dimensions thatdefine the airfoil characteristics of the part. The shroud section 40 isengaged with and held by an annular housing member (not shown). Aplurality of interlocking blades are thus assembled with the housingmember to form a disc. In the operating gas turbine engine the assembleddiscs, which are rotating parts, determine the path of the intake,combustion and exhaust gasses that flow through the engine.

FIG. 23(b) shows two adjacent blades 38 of an assembled disc. The bladesare held in the housing member (not shown) such that surfaces 44 of eachshroud section 40 contacts corresponding surfaces 44 of adjacentshrouds. These contact surfaces 44 are subjected to wearing forcesduring the operation of the gas turbine engine. As an assembled disc ofblades rotates, the individual adjacent blades 38 may chatter againsteach other, causing wear to occur at the contact surfaces 46 of theshroud sections 38. This chattering results in constant hammering at thecontact surfaces 44 of the interlocking blades 38. Excessive wear in thearea of the contact surfaces 44 can have detrimental consequences on theoperation of the gas turbine engine, and thus is an area of concern.

To combat the excessive wear in the area of the contact surfaces of theshrouds, it has been conventional practice to apply a hard facingmaterial to the shroud in the location of the contact surfaces. FIG.23(a) shows a typical location for the application of a hard facingmaterial 46. The hard facing material is applied to the shroud by, forexample, manual tig welding or laser welding.

As disclosed in co-owned US patent applications and as disclosed herein,applicants have invented methods for creating a diffusion bonded coatingon the surface of a workpiece, such as a turbine engine airfoil part.These co-owned applications include a application Ser. No. 10/638,192,which is a Continuation-in-Part of application Ser. No. 10/423,722,filed Apr. 28, 2003, which is a Continuation-in-Part of application Ser.No. 10/241,854, filed Sep. 13, 2002, which is a Continuation-in-Part ofapplication Ser. No. 09/505,803, filed Feb. 17, 2000, which is aContinuation-in-Part of application Ser. No. 09/143,643, filed Sep. 3,1998, now U.S. Pat. No. 6,049,978, which is a Continuation-in-Part ofapplication Ser. No. 08/993,116, now U.S. Pat. No. 5,956,845, which isthe utility patent application of a U.S. provisional application Ser.No. 60/033,858, filed Dec. 23, 1996; and relates to an inventiondisclosed in an Invention Disclosure Document accepted under theDisclosure Document program on or about Nov. 5, 1996 and assignedDisclosure Document No. 407616.

In accordance with the invention described in applicants co-pending U.S.patent application, Ser. No. 10/021,107, which is incorporated byreference herein, a cobalt-based alloy is provided that is particularlyuseful as a hard facing material for gas turbine engine components, suchas the shrouds of a gas turbine engine blade. The alloy compositions asdescribed in this co-pending application have a relatively smalllanthanum addition and relatively large carbon content and provideremarkable oxidation resistance and wear resistance at hightemperatures. Importantly, the inventive alloy composition has asuitable combination of ductility and wear resistance at hightemperatures to be effective as a hard face material for limiting theeffects of chattering of blades during the operation of a gas turbineengine. Accordingly, the inventive alloy has a suitable combination ofductility, oxidation resistance and wear resistance and thus representsan improved hard facing material for the blade components of gas turbineengine.

A conventional hard facing material for use on the blade of gas turbineengines consists of an alloy containing chromium, tungsten, nickel andcobalt. U.S. Pat. No. 3,265,434, issued to Baldwin, teaches an alloy forhigh temperature use containing chromium, tungsten, nickel and cobalt.Baldwin specifically teaches an alloy with improved short time tensilestrength at 180° F., wherein the ratio of cobalt to chromium is alwaysat least 1.4:1. Baldwin further teaches that an alloy with optimumcharacteristics, from the standpoint of a combination of ductility(freedom from brittleness), and wear resistance, were obtained with anickel content in the range of 4 to 6%. The composition taught byBaldwin has a short time tensile strength at 1800° F. of 48,000 p.s.i.

U.S. Pat. No. 3,582,320, issued to Herchenroeder, teaches a cobalt basealloy having superior oxidation and wear resistance. Herchenroederteaches that a relatively small lanthanum addition and a relativelylarge carbon content provides remarkable oxidation resistance and wearresistant properties at high temperatures. The composition taught byHerchenroeder has an ultimate tensile strength of 15,700 p.s.i.

U.S. Pat. No. 3,947,269, issued to Prasse et al., teaches aboron-hardened tungsten facing alloy used as a facing or coating forbase material, and in particular as a piston ring facing. The alloytaught be Prasse et al. is applied as a metal powder that is melted andsprayed upon a workpiece, such as a piston ring of a high compressioncombustion engine.

U.S. Pat. No. 4,822,248, issued to Wertz et al., teaches a method ofrebuilding a shroud of a turbine blade. A wear resistant overlay isformed using a plasma torch. A powdered metal is applied to the notch ofa shroud and a plasma transferred arc is generated at low amperagesufficient to melt and cast the powdered metal while holding the heatimparted to the shroud to a minimum.

To be effective for use in the demanding environments subjected to theblades in an operating gas turbine engine, a hard facing material musthave superior oxidation and wear resistance at elevated temperatures.Further, the hard facing material must have a suitable degree ofductility to withstand the constant hammering caused by chatteringblades. As shown in FIG. 23(b), the contact surfaces 44 of the shroudsare subjected to wearing forces during the operation of the gas turbineengine. As an assembled disc of blades rotates, the individual adjacentblades 38 may chatter against each other, causing wear to occur at thecontact surfaces 44 of the shroud sections 40. This chattering resultsin constant hammering at the contact surfaces 44 of the interlockingblades 38. Excessive wear in the area of the contact surfaces 44 causedby chipping and flaking can have detrimental consequences on theoperation of the gas turbine engine. Accordingly, there is a need for amethod for forming a wear-resistant hard-face contact area on aworkpiece, such as a gas turbine engine part that is subjected tofailure due to chipping or flaking.

SUMMARY OF THE INVENTION

The present invention overcomes the drawbacks of the conventional. It isan object of the present invention to provide a method for correctingthe dimensional characteristics of a cast article. It is another objectof the present invention to provide a method for applying a protectivecoating to a metal article. It is another object of the presentinvention to provide a method for repairing a turbine engine airfoilpart.

In accordance with the present invention, a method of correcting thedimensional characteristics of a cast article is provided. Thedimensional differences are determined between pre-repair cast articledimensions and desired post repair cast article dimensions to correct acasting defect in the article. The determination may be made bydetermining the location and approximate volume of a void in the surfaceof the article. The determination may also be made by determining anamount of buildup volume required to make at least a portion of thesurface of the cast article built up to the desired post repairdimensions. The article is coated in at least an area of the castingdefect with a high-density coating material. Depending on the coatingprocess, the coating can be formed in a vacuum, inert atmosphere orunder ambient conditions. How ever it is applied, in accordance with theinvention, the coating must be capable of forming a diffusion boundarybetween the coating material and the article. A hot isostatic heattreatment process is performed to form the diffusion boundary betweenthe coating material and the article.

Depending on the type of casting defect, material in an area of thecasting defect may be removed before the step of coating the article.For example, if the casting defect is an inclusion of an undesiredcomposition, such as an oxide or dirt particle, the inclusion and someof the base article material can be removed by a machining or otheroperation. The area of the casting defect is enlarged, and may becontoured to create a better surface for holding the coating material.The casting defect may be caused, for example, by at least one of aninclusion at a surface of the article, an air bubble at the surface ofthe article, undercasting, a void and shrinkage. A sintering heattreatment can be performed before the step of performing the hotisostatic heat treatment to limit the occurrence of bubbles on thesurface of the coating material after an isostatic heat treatment. Thesintering heat treatment may be performed at a temperature substantiallythe same as the temperature of the hot isostatic heat treatment.

In accordance with the present invention, the coating material maycomprise an alloy with substantially no oxide forming constituents so asto avoid the formation of oxide inclusions in the coating material. Inthis case, the coating material may be applied using a coating processthat is effective to create a coating on the surface of the article thatwill be diffusion bonded to the article after the hot isostatic heattreatment.

In accordance with the present invention, a method is provided forapplying a protective coating to a metal article. A metal article isprovided and coated with a high-density coating material capable offorming a diffusion boundary between the coating material and thearticle. In accordance with this aspect of the invention, the coatingmaterial comprises an alloy with substantially no oxide formingconstituents so as to avoid the formation of oxide inclusions in thecoating material. Applicants have discovered that the oxides in thecoating may form crack initiation sites, and cracks formed due to theoxides may propagate through the diffusion boundary and into the articlesubstrate. By limiting the formation of oxides in the coating, thesecrack initiation sites are reduced or eliminated, thereby enabling thecoating material to act as a protective coating. Depending on thecoating process, the coating may be applied in a vacuum, under an inertatmosphere or under ambient conditions. In the case of ambientconditions in which oxygen may be present, the coating material may becomposed of constituents that substantially avoid the formation of oxideparticles, even when oxygen is present. Stated otherwise, the coatingmaterial has a chemistry that does not result in crack producingelements, such as oxides, located in the coating and in the diffusionboundary between the coating and the substrate.

The hot isostatic heat treatment process is performed to form thediffusion boundary between the coating material and the article. Thus,in accordance with this aspect of the invention, the substantially oxidefree coating and the diffusion boundary provide a protective coating toprotect the article from damage.

A sintering heat treatment can be performed before the step ofperforming the hot isostatic heat treatment to limit the occurrence ofbubbles on the surface of the coating material after an isostatic heattreatment. The sintering heat treatment may be performed at atemperature substantially the same as the temperature of the hotisostatic heat treatment. The sintering heat treatment increases theproduction yield by significantly reducing the formation of bubbles onthe surface of the coating due to the hot isostatic heat treatment, etc.

In accordance with another aspect of the invention, a method is providedfor repairing a turbine engine airfoil part. The dimensional differencesare determined between pre-repair airfoil dimensions of a turbine engineairfoil part substrate and desired post repair airfoil dimensions of theturbine engine airfoil part substrate. The pre-repair airfoil dimensionshaving different airfoil characteristics than the post-repair airfoildimensions. The turbine engine airfoil part being comprised of a metalalloy. The engine airfoil part is coated with a coating capable offorming a diffusion boundary with the turbine engine airfoil partsubstrate. The coating material comprises an alloy with substantially nooxide forming constituents so as to avoid the formation of oxideinclusions in the coating material. A hot isostatic heat treatmentprocess is performed to obtain a post-repair turbine engine airfoil parthaving the desired post-repair dimensions and having a substantiallyoxide free coating and diffusion bonding between the coating materialand the turbine engine airfoil part substrate. The substantiallyoxide-free coating provides a protective coating to protect the articlefrom damage. A sintering heat treatment can be performed before the stepof performing the hot isostatic heat treatment to limit the occurrencebubbles on the surface of the coating material after an isostatic heattreatment. The sintering heat treatment may be performed at atemperature substantially the same as the temperature of the hotisostatic heat treatment.

In accordance with the present invention, a method is provided forforming a diffusion coating on the surface of a workpiece. A workpiecesubstrate is provided. A coating is formed on at least one of theselected portions of the workpiece substrate. The coating material iscapable of forming a diffusion bond with the workpiece substrate. Thediffusion bond is a metallurgical bond between the workpiece and thecoating that does not have an interface boundary. This diffusion bondcreates a secure attachment between the coating and the substrate, muchstronger than the mechanical bond that is originally formed between thecoating and the substrate. A sintering heat treatment is first performedto expel trapped gas from the coating material. Applicant has found thatthe entrapped gas is problematic because it results in a weaker, bubbledsurface with an inconsistent diffusion bond between the coating and thesubstrate. The sintering heat treatment removes the entrapped gas andprevents outgassing of the trapped gas during a hot isostatic pressingtreatment. This preventive treatment has been experimentally proven togreatly reduces the formation of bubbles on the surface of the coatedworkpiece after the hot isostatic pressing treatment. After theentrapped gas is removed by the sintering heat treatment, the hotisostatic pressing treatment is then performed to drive the coatingmaterial into the workpiece substrate. The hot isostatic pressingtreatment results in the formation of the diffusion bond so that themetallurgical bond between the workpiece and the coating is formed.

A method of correcting defects in a metal workpeice. A location of adefect in a workpiece is determined. The defect comprising a void or aninclusion in a workpiece substrate. The workpiece substrate is comprisedof a metal alloy. Material of the workpiece substrate at the location ofthe defect is removed to form a cleaned area in the workpiece substrate.The cleaned area in the workpiece substrate is coated with ahigh-density coating. A sintering heat treatment is performed on thecoated workpiece substrate to remove entrapped gas from the coatingmaterial prior to a step of hot isostatic pressing treatment. Then, hotisostatic pressing treatment is performed on the coated workpiece toproduce diffusion bonding between the workpiece substrate and thehigh-density coating. The material can be removed by techniques such assandblasting or grinding. A high-density coating process such ashyper-velocity oxy-fuel thermal spray process or a detonation gunprocess is used to apply the high-density coating to the substrate atthe location of the cleaned area. The high-density coating may have thesame metal alloy composition as the metal alloy substrate. The metalalloy substrate may comprise a nickel or cobalt-based superalloy, andthe high-density coating may have the same nickel or cobalt-based superalloy composition as the metal alloy substrate.

The workpiece substrate is prepared for a high-density coating process.The preparation may include cleaning, blasting, machining, masking orother like operations. Once the workpiece substrate has been prepared, ahigh-density coating process is performed to coat the workpiecesubstrate. The coating material is built-up to a thickness that iseffective to obtain desired finished dimensions after performing a hotisostatic pressing treatment (described below). The high-density coatingprocess may comprise performing a hyper velocity oxy-fuel thermal sprayprocess. In the case of HVOF, a fuel gas and oxygen are used to create acombustion flame at 2500 to 3100° C. The combustion takes place at avery high chamber pressure and a supersonic gas stream forces thecoating material through a small-diameter barrel at very high particlevelocities. The HVOF process results in extremely dense, well-bondedcoatings. Typically, HVOF coatings can be formed nearly 100% dense, withat a porosity of about 0.5%. The high particle velocities obtained usingthe HVOF process results in relatively better bonding between thecoating material and the substrate, as compared with other coatingmethods such as the conventional plasma spray method or the chemicalvapor deposition method. However, the HVOF process forms a bond betweenthe coating material and the substrate that occurs primarily throughmechanical adhesion at a bonding interface. As will be described below,in accordance with the present invention this mechanical bond isconverted to a metallurgical bond by creating a diffusion bond betweenthe coating material and the workpiece substrate. This diffusion bonddoes not have the interface boundary which is usually the site offailure.

The diffusion bond is created by subjecting the coated workpiecesubstrate (or, in the case of the inventive repair method, the coatedairfoil part) to a hot isostatic pressing (HIP) treatment. Theappropriate hot isostatic pressing treatment parameters are selecteddepending on the coating, the workpiece substrate and the finalattributes that are desired. The hot isostatic pressing treatment isperformed on the coated workpiece substrate to obtain a metal producthaving the desired finished dimensions and diffusion bonding between thecoating material and the workpiece substrate.

HIP treatment is conventionally used in the densification of cast metalcomponents and as a diffusion bonding technique for consolidating powdermetals. In the HIP treatment process, a part to be treated is raised toa high temperature and isostatic pressure. Typically, the part is heatedto 0.6-0.8 times the melting point of the material comprising the part,and subjected to pressures on the order of 0.2 to 0.5 times the yieldstrength of the material. Pressurization is achieved by pumping an inertgas, such as Argon, into a pressure vessel. Within the pressure vesselis a high temperature furnace, which heats the gas to the desiredtemperature. The temperature and pressure is held for a set length oftime, and then the gas is cooled and vented.

In accordance with the present invention, the HIP treatment process isperformed on a HVOF coated substrate to convert the adhesion bond, whichis merely a mechanical bond, to a diffusion bond, which is ametallurgical bond. In accordance with the present invention, an HVOFcoating process is used to apply the coating material having sufficientdensity to effectively undergo the densification changes that occurduring the HIP process. After the HVOF spray material is applied, asintering heat treatment process can be performed to further densify thecoating to prevent gas entrapment of the coating material and/or thediffusion bonding area during the hot isostatic pressing process. If thecoating material and the workpiece substrate are comprised of the samemetal composition, then the diffusion bonding results in a particularlyseamless transition between the substrate and the coating.

The inventive method can be used for forming a metal product having awear resistant surface. This method can be employed to produce, forexample, a long lasting cutting tool from a relatively inexpensivecutting tool substrate. In accordance with this aspect of the invention,a workpiece substrate is formed to near-finished dimensions. Ahigh-density coating process, such as a hyper velocity oxy-fuel thermalspray process, is performed to coat the workpiece substrate with a wearresistant coating material. The coating material is built-up to athickness that is effective to obtain desired finished dimensions afterperforming a hot isostatic pressing treatment. A sintering heattreatment step may be performed to improve the density of the coatingmaterial and prevent gas entrapment during the hot isostatic pressingtreatment. The hot isostatic pressing treatment is performed on thecoated workpiece substrate to obtain a metal product having the desiredfinished dimensions and diffusion bonding between the coating materialand the workpiece substrate.

The inventive method can also be used for forming a cast metal product.This method can be employed to produce, for example, a cast part havinga hard and/or smooth surface. In accordance with the present invention,a part is cast to dimensions to less than the finished dimensions, or acast part is machined to less than the finished dimensions. The castpart is then coated using the HVOF coating method as described herein.The HVOF coating is applied to a thickness sufficient to bring the partto its finished dimensions. The HVOF coated, cast part is then HIPtreated as described herein to obtain a finished part having desireddimensions and surface characteristics.

In accordance with this aspect of the invention, a cast metal workpieceis provided. The cast metal workpiece may be formed from anyconventional casting method such as: investment, sand and resin shellcasting.

The cast metal workpiece is machined, if necessary, to near-finisheddimensions. A high-density coating process, such as a hyper velocityoxy-fuel thermal spray process (HVOF), is performed to coat theworkpiece substrate with a coating material. The coating material isbuilt-up to a thickness effective to obtain desired finished dimensionsafter performing a hot isostatic pressing treatment. A sintering heattreatment step may be performed to improve the density of the coatingmaterial and prevent gas entrapment during the hot isostatic pressingtreatment. The hot isostatic pressing treatment is performed on thecoated workpiece substrate to obtain a metal product having the desiredfinished dimensions and diffusion bonding between the coating materialand the workpiece substrate.

In accordance with another aspect of the present invention, thereclassification of a gas turbine engine airfoil part is obtained. Thedimensional differences between pre-reclassified dimensions of thebuttresses of a turbine engine airfoil part and desiredpost-reclassified dimensions of the buttresses are determined. That is,the change in shape of the inner buttress and outer buttress necessaryto obtain a desired angular relationship between the airfoil section andthe buttresses is determined. Build-up thickness of coating materialrequired to obtain the desired post-reclassified dimensions of thebuttresses is determined. A high-density coating process, such as HVOF,is used to coat the buttresses of the turbine engine airfoil part with acoating material. The portions of the part that are not to be built up,such as the airfoil section and parts of the buttresses, may be maskedbefore applying the high-density coating. Also, some of the coatedsurfaces of the part may need to be built up more than others. Thecoating material is applied to the determined build-up thickness ofcoating material effective to obtain the desired post-reclassificationdimensions after performing a hot isostatic pressing treatment, andafter the selective removal of some of the original buttress materialand some of the built up coating material. A sintering heat treatmentmay be performed before the hot isostatic pressing treatment.

As discussed herein, the coating material comprises a metal alloycapable of forming a diffusion bond with the substrate of the turbineengine airfoil part. After the coating material is applied, thesintering heat treatment process may be performed to prevent gasentrapment of the coating material and/or the diffusion bonding areaduring the hot isostatic pressing process. Then, the hot isostaticpressing (HIP) process is performed so that the buttresses of theturbine engine airfoil part have a robust diffusion bonding between thecoating material and the original material of the buttresses. Havingbuilt up the appropriate dimensions of the inner buttress and outerbuttress, the reclassification of the part is obtained by selectivelyremoving the original buttress material and, if necessary, some of thebuilt up material until the angular relationship between the airfoilsection and the inner and outer buttresses is obtained. The material canbe removed through milling, grinding, or other suitable and well knownmachining operations. Further, to facilitate obtaining the correctdimensions the centerline position of the airfoil part can be locatedand held by mounting the part in a suitable holding fixture whenmachining the buttresses.

The fixture may be so constructed so that a vane that has at least aminimum amount of material built up on its buttresses can be machinedand reclassified. In this case, it may not be necessary to determine thedimensional differences or the required build-up thickness. Rather, theinventive high density coating and HIPing process (and, if neededsintering) can be performed to build up at least the minimum amount ofmaterial diffusion bonded to the buttresses. Then, the vane is placed inthe fixture and the excess material (both original buttress material andthe built-up material) is machined until the buttresses have beenreshaped and the vane reclassified as intended.

In accordance with another aspect of the present invention, a method isprovided for forming a wear-resistant hardfaced contact area on theshroud section of a gas turbine engine blade. A predetermined contactarea of a shroud section of a gas turbine engine blade is selectivelycoated with a high-density hardface coating material. The hardfacecoating material is capable of forming a diffusion boundary between thehardface coating material and the shroud section. A hot isostatic heattreatment process is performed to form the diffusion boundary betweenthe hardface coating material and the shroud section to form awear-resistant hardfaced contact area diffusion bonded to the shroudsection.

Depending on the coating process, and the necessity for doing so, thepredetermined contact area can be masked off before the step ofselectively coating. A sintering heat treatment can be perfomed beforethe step of performing the hot isostatic heat treatment to limit theoccurrence of bubbles on the surface of the hardface coating materialafter the isostatic heat treatment step. The sintering heat treatmentmay be performed at a temperature substantially the same as thetemperature of the hot isostatic heat treatment. The hardface coatingmaterial may comprise an alloy with substantially no oxide formingconstituents so as to avoid the formation of oxide inclusions in thecoating material.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1(a) is a flow chart showing the steps of the inventive method forrepairing a gas turbine engine airfoil part;

FIG. 1(b) is a flow chart showing the steps of the inventive method offorming metal products and metal components having a wear resistantcoating; FIG. 1(c) is a flow chart showing the steps of the inventivemethod for correcting defects in a workpiece;

FIG. 2(a) is a schematic view of a tool substrate provided in accordancewith the inventive method of forming metal components having a wearresistant coating;

FIG. 2(b) is a schematic view of the tool substrate having a wearresistant coating applied using an HVOF thermal spray process inaccordance with the inventive method of treating metal components havinga wear resistant coating;

FIG. 2(c) is a schematic view of the HVOF spray coated tool substrateundergoing a HIP treatment process in a HIP vessel in accordance withthe inventive method of forming metal components having a wear resistantcoating;

FIG. 2(d) is a schematic view of the final HVOF spray coated and HIPtreated tool having a wear resistant coating layer diffusion bonded tothe tool substrate in accordance with the inventive method of formingmetal components having a wear resistant coating;

FIG. 3(a) is a schematic perspective view of a cast metal componentundergoing a machining operation in accordance with the inventive methodof forming a metal product;

FIG. 3(b) is a schematic perspective view of the machined cast metalcomponent in accordance with the inventive method of forming a metalproduct;

FIG. 3(c) is a schematic perspective view of the machined cast metalcomponent having a coating applied using an HVOF thermal spray processin accordance with the inventive method of forming a metal product;

FIG. 3(d) is a schematic perspective view of the HVOF spray coatedmachined cast metal component undergoing a HIP treatment process in aHIP vessel in accordance with the inventive method of forming a metalproduct;

FIG. 3(e) is a schematic perspective view of the final HVOF spray coatedand HIP treated machined cast metal product having a coating layerdiffusion bonded to the machined cast metal component in accordance withthe inventive method of forming a metal product;

FIG. 4 is a flow chart showing the steps of the inventive method ofrepairing a turbine engine part;

FIG. 5(a) is a schematic side view of a worn turbine engine part beforeundergoing the inventive method of repairing a turbine engine part;

FIG. 5(b) is a schematic cross-sectional view of the worn turbine enginepart before undergoing the inventive method of repairing a turbineengine part;

FIG. 6(a) is a schematic side view of the worn turbine engine partshowing the worn areas to be repaired using the inventive method ofrepairing a turbine engine part;

FIG. 6(b) is a schematic cross-sectional view of the worn turbine enginepart showing the worn areas to be repaired using the inventive method ofrepairing a turbine engine part;

FIG. 7(a) is a schematic side view of the worn turbine engine partshowing the worn areas filled in with similar weld material inaccordance with the inventive method of repairing a turbine engine part;

FIG. 7(b) is a schematic cross-sectional view of the worn turbine enginepart showing the worn areas filled in with similar weld material inaccordance with the inventive method of repairing a turbine engine part;

FIG. 8(a) is a schematic side view of the welded turbine engine partshowing areas to be built up with similar coating material using an HVOFspray coating process in accordance with the inventive method ofrepairing a turbine engine part;

FIG. 8(b) is a schematic cross-sectional view of the welded turbineengine part showing areas to be built up with similar coating materialusing an HVOF spray coating process in accordance with the inventivemethod of repairing a turbine engine part;

FIG. 9(a) is a schematic side view of the HVOF built up, welded turbineengine part showing an area masked before performing the HVOF spraycoating process in accordance with the inventive method of repairing aturbine engine part;

FIG. 9(b) is a schematic cross-sectional view of the HVOF built up,welded turbine engine part in accordance with the inventive method ofrepairing a turbine engine part;

FIG. 10 is a schematic view of the HVOF built up, welded turbine enginepart undergoing a HIP treatment process in a HIP vessel in accordancewith the inventive method of repairing a turbine engine part;

FIG. 11(a) is a schematic side view of the final HVOF spray coated andHIP repaired turbine engine part having a similar metal coating layerdiffusion bonded to the original parent substrate and welded portions inaccordance with the inventive method of repairing a turbine engine part;

FIG. 11(b) is a schematic cross-sectional view of the final HVOF spraycoated and HIP repaired turbine engine part having a similar metalcoating layer diffusion bonded to the original parent substrate andwelded portions in accordance with the inventive method of repairing aturbine engine part;

FIG. 12(a) is a side view of a prior art tool bit coated with a wearresistant coating;

FIG. 12(b) is a side view of a prior art tool bit having a fixed wearresistant cutting tip;

FIG. 13 is a flow chart showing the steps of the inventive method forreclassifying a gas turbine engine airfoil part;

FIG. 14(a) is a front view of a vane from a gas turbine engine showingthe airfoil section, the outer buttress and the inner buttress;

FIG. 14(b) is a partial top view of the vane shown in FIG. 14(a) showingthe outer buttress and angle α indicating the angular relationshipbetween the airfoil and the outer buttress;

FIG. 14(c) is a partial bottom view of the vane shown in FIG. 14(a)showing the inner buttress and angle α′ indicating the angularrelationship between the airfoil and the inner buttress;

FIG. 14(d) is a partial left-side view of the vane shown in FIG. 14(a)showing the leading edge foot of the inner buttress and the outer footfront face of a buttress rail of the outer buttress;

FIG. 14(e) is a partial right-side view of the vane shown in FIG. 14(a)showing the trailing edge foot of the inner diameter buttress and theother buttress rail of the outer diameter buttress;

FIG. 15(a) is a flowchart showing the steps of the inventive method forrepairing a workpiece with an electroplated coating diffusion bonded tothe workpiece;

FIG. 15(b) is a flow chart showing the steps of the inventive method forrepairing a gas turbine engine airfoil part with an electroplatedcoating diffusion bonded to the airfoil substrate;

FIG. 15(c) is a flow chart showing the steps of the inventive method forcorrecting defects in a workpiece with an electroplated coatingdiffusion bonded to the workpiece;

FIG. 15(d) is a flow chart showing the steps of the inventive method forreclassifying a gas turbine engine airfoil part with an electroplatedcoating diffusion bonded to the airfoil part;

FIG. 16(a) shows an airfoil part prepared for electroplating;

FIG. 16(b) shows the prepared airfoil part being electroplated;

FIG. 16(c) shows the electroplated airfoil part undergoing a sinteringheat treatment;

FIG. 16(d) shows the sintered electroplated airfoil part undergoing ahot isostatic heat treatment;

FIG. 16(e) shows the finished airfoil part having a diffusion bondbetween the electroplated areas and the airfoil substrate;

FIG. 17 illustrates the steps of correcting the dimensionalcharacteristics of a cast article;

FIG. 18 is a flow chart showing the steps of the inventive method ofcorrecting the dimensional characteristics of a cast article;

FIG. 19 schematically illustrates a coated substrate wherein the coatingmaterial is diffusion bonded to the substrate and includes an oxideinclusion;

FIG. 20 schematically illustrates the coated substrate shown in FIG. 19wherein a crack is forming at the site of the oxide inclusion;

FIG. 21 schematically illustrates the coated substrate shown in FIG. 19wherein the crack formed at the site of the oxide inclusion propagatesthrough the diffusion boundary and into the substrate;

FIG. 22(a) is a schematic perspective view of a cast turbine engineairfoil part showing a casting defect;

FIG. 22(b) is a schematic perspective view of the cast turbine engineairfoil part having the area of the casting defect being machined;

FIG. 22(c) is a schematic perspective view of the cast turbine engineairfoil part after the area of the casting defect has been machined;

FIG. 23(d) is a schematic perspective view of the cast turbine engineairfoil part having the area of the casting defect being filled with acoating material;

FIG. 22(e) is a schematic perspective view of the coated cast turbineengine airfoil part being subjected to a hot isostatic pressingtreatment;

FIG. 22(f) is a schematic perspective view of the repaired cast turbineengine airfoil part;

FIG. 23(a) is a cut-away view of a gas tubine engine blade showing theshroud portion afixed to the airfoil portion of the blade, and showingthe location of an applied wear resistant hard facing material to thecontact surface of the blade;

FIG. 23(b) shows two adjacent blades of an assembled disc showing thecontact between the shrouds of the blades; and

FIG. 23(c) is a flow chart illustrating the steps of the inventivemethod for forming a wear-resistant hard-face contact area on aworkpiece, such as a gas turbine engine part.

DETAILED DESCRIPTION OF THE INVENTION

For purposes of promoting an understanding of the principles of theinvention, reference will now be made to the embodiments illustrated inthe drawings and specific language will be used to describe the same. Itwill nevertheless be understood that no limitation of the scope of theinvention is thereby intended, there being contemplated such alterationsand modifications of the illustrated device, and such furtherapplications of the principles of the invention as disclosed herein, aswould normally occur to one skilled in the art to which the inventionpertains.

Referring to FIG. 1(a), in accordance with the present invention, thedimensional differences between pre-repaired dimensions of a turbineengine airfoil part and desired post-repair dimensions of the turbineengine airfoil part are determined (Step One-B). The turbine engineairfoil part has a substrate comprised of a superalloy. A build-upthickness of coating material required to obtain the desired post-repairdimensions of the turbine engine airfoil part is determined (Step Two).A high-density coating process, such as HVOF, is used to coat theturbine engine airfoil part with a coating material to the determinedbuild-up thickness of coating material effective to obtain the desiredpost-repair dimensions after performing a sintering heat treatment and ahot isostatic pressing treatment (Step Three). The coating materialcomprises a metal alloy capable of forming a diffusion bond with thesubstrate of the turbine engine airfoil part. After the coating materialis applied, a sintering heat treatment process is performed to preventgas entrapment of the coating material and/or the diffusion bonding areaduring the hot isostatic pressing process (Step Four). Then, the hotisostatic pressing process is performed to obtain a post-repair turbineengine airfoil part having the desired post-repair dimensions and havingdiffusion bonding between the coating material and the turbine engineairfoil substrate (Step Five).

In accordance with the present invention, a protective coating must befirst removed from the turbine engine airfoil part prior to performingthe high-density coating process (Step One-A). After performing the hotisostatic pressing process, a protective coating may be re-applied (StepSix). In this case, the build-up thickness may be determined in Step Twoto take into consideration the additional thickness of the post-repairedpart due to the addition of the protective coating.

Typically, this protective coating is present on an airfoil part toprotect it from the hot corrosive environment it experiences duringservice. This protective coating must be removed during the inspectionand/or repair process. After undergoing a number of inspection and/orrepair cycles, the airfoil part was conventionally discarded simplybecause the airfoil dimensions of the part were too deformed for thepart to be usable. However, in accordance with the present inventiverepair method, the airfoil dimensions are restored and a robust repairedairfoil part is obtained

In the typical application of the inventive method, the metal alloysubstrate of the turbine engine airfoil part will comprise a nickel orcobalt-base superalloy. The step of performing the high-density coatingprocess (Step Three) may thus include performing a high-density coatingprocess such as a hyper velocity oxy-fuel thermal spray process or adetonation gun process to apply a high-density coating having the samenickel or cobalt-base superalloy composition as the metal alloysubstrate.

In an embodiment of the invention in which the coating material and thesubstrate alloy comprise INCO713C nickel or cobalt-base superalloy, thesintering heat treatment (Step Four) comprises sintering at atemperature at or about 2150 degrees F. for about 2 hours, which hasbeen found to effectively prevent gas entrapment of the appliedhigh-density coating during the hot isostatic pressing process. Therange at which the sintering heat treatment may be performed is about1900 to 2300 degrees F. In the case of the nickel or cobalt-basesuperalloy substrate, an effective hot isostatic pressing treatment(Step Five) can be performed at a temperature of about 2200 F. in about15 KSI argon for about 4 hours. The inventive process may be used withalloys of other metals, such as titanium or aluminum. The parameters ofthe hot isostatic pressing treatment typically call for heating theengine part to a temperature that is substantially 80% of the meltingpoint of the metal alloy; and pressurizing the engine part to a pressuresubstantially between 20 and 50 percent of the yield strength of themetal alloy in an inert gas atmosphere.

The dimensional differences between the pre-repaired dimensions of theturbine engine airfoil part and the desired post-repair dimensions ofthe turbine engine airfoil part are measured from at least one of thecordal and length dimensions of the airfoil part (Step One-B). Byperforming the inventive method for repairing a gas turbine engineairfoil part, the post-repair dimensions are equal to the dimensionsnecessary for effectively returning the part to active service. Theobtained diffusion bonding between the coating material and thesubstrate ensures that the repaired airfoil part is robust enough towithstand the highly demanding environmental conditions present in anoperating gas turbine engine. Thus, the present invention offerssubstantial cost savings over having to replace a turbine gas engineairfoil part which otherwise might have been discarded.

The present invention can be used as a process for restoring criticalgas path area dimensions in cast nickel or cobalt-base superalloy vanecomponents. These dimensions may become altered due to erosion orparticle strikes during the service life of the part, and/or may becomealtered during an inspection or repair process wherein a protectivecoating is stripped from the part.

The inventive process, referred to herein as “recast”, briefly consistsof applying a pre-alloyed metal powder, compositionally identical to thesuperalloy used in the original manufacture of the vane being repaired,directly on dimensionally discrepant surfaces, densifying the metalpowder coating, and causing it to bond to the affected surface.

More specifically, in the preferred embodiment of the inventioncandidate recast surfaces are abrasively clean, thermal sprayed usinghigh velocity oxy fuel processes (HVOF), sintered, and hot isostaticallypressed (HIPed).

Thermal spray metal powders, produced by a vacuum/inert gas atomizationprocesses, are applied directly to the dimensionally discrepant surfacesof a turbine engine airfoil part using robotic HVOF processes carefullycontrolled to produce dense coatings while minimizing thermal gradientsand oxidative solute losses.

Properly applied HVOF coatings are dense but sometimes containinterconnected micropores. In accordance with the present invention,such “porous” HVOF coatings are more fully densified by sintering andsubsequently diffusion-bonded to substrate surfaces by HIPing attemperatures and pressures commensurate with the nickel or cobalt-basealloy under consideration.

Recast surfaces are compositionally identical to, but microstructurallydifferent from, original or “as-cast” substrates. As-cast substrates aredefined herein as a substrate formed by a conventional casting process,such as the lost wax or investment casting process described above. Themicrostructures of cast nickel or cobalt-base superalloy substratematerials such as used in the manufacture of gas turbine vanes generallyconsist of a relatively large amount of an intermetallic precipitatereferred to as “gamma prime” within, and networks of carbides andborides within and around, large “gamma” matrix grains. The amount andmorphology of gamma prime, carbides, and borides are determined bycomposition, processing history, and heat treatment.

Recast microstructures similarly consist of gamma prime, carbides, andborides precipitated in and around gamma matrix grains; but, recastmatrix grains are considerably smaller than as-cast grains. Recast gammaprime, carbide and boride precipitates are similarly finer than as-cast.In addition, some of the more reactive solutes (e.g., aluminum) in thethermal spray powders oxidize during the HVOF spray process to formoxide particles which become randomly dispersed in the recast deposit.

Articles repaired by recast are best described as bimetallic compositescomprised of recast coatings bonded to as-cast substrates. Themechanical properties of such repaired articles vary depending on therelative volume fraction of the recast coating, the specific alloy(s)under consideration, and processing history.

Example of Recast INCO713C/cast INCO713C Composite Mechanical PropertiesObtained in Accordance with the Present Invention

Representative tensile and stress-rupture properties of recastINCO713C/cast INCO713C composite test specimens were measured to morefully elucidate the recast process.

INCO713C was selected as the base nickel or cobalt-base superalloy formeasurement because it is specified by a large number of enginemanufactures for gas turbine component applications, and isbill-of-material for JT8D second-stage vanes, a candidate component forthe inventive recast repair method.

Near cast-to-size INCO713C test bars were machined into ASTMproportioned mechanical test specimens with tapered (approximately threepercent) gauge lengths. The average minimum gauge length diameter was0.2137 inches.

The machined test specimens were grit-blasted with silicon carbide,ultrasonically cleaned, and robotically sprayed with INCO713C powderusing Diamond Jet HVOF processes. The composition of the INCO713C powderused in these evaluations is shown in Table I. TABLE I CertifiedCompositions of INCO713C Atomized Powder and Cast-To-Size Test BarsCast-To-Size Test Bars Element EMS 55079 Atomized Powder (Heat #8616)Nickel Balance Balance Balance Chromium 11.0 to 13.0 13.6 13.67 Aluminum5.5 to 6.5 5.86 5.61 Molybdenum 3.8 to 5.2 4.39 4.06 Columbium 1.5 to2.5 2.1 2.08 Titanium 0.4 to 1.0 0.9 0.84 Zirconium 0.05 to 0.15 0.070.05 Carbon 0.05 to 0.07 0.1 0.13 Boron 0.005 to 0.015 0.01 0.008 Cobalt1.00 max. <0.01 <0.05 Silicon 0.50 max. 0.09 <0.05 Copper 0.05 max. 0.04<0.05 Iron 0.25 max. 0.18 <0.05 Manganese 0.25 max. 0.01 <0.05 Sulfur0.015 max. 0.002 <0.05 Phosphorus 0.015 max.

Sufficient HVOF coating was applied to increase the composite specimengauge length diameter to approximately 0.250 inches. The sprayed testbars were then sintered at 2150 F for 2 hours in vacuum, HIPed at 2200 Fin 15 KSI argon for 4 hours in a standard commercial HIP toll cycle, andtested for room temperature tensile and elevated-temperaturestress-rupture.

The composite test specimens used for these measurements were nominallycomprised of 28 percent recast INCO713C and 72 percent as-cast INCO713C.The recast INCO713C percentage varied, however, from 25.5 to 30.9percent depending on precise machined and sprayed specimen dimensions.

Mechanical Properties:

The room temperature tensile and 1800 F stress-rupture properties of theas-cast INCO713C core material used in these measurements are summarizedin Table II. TABLE II INCO713C Heat # 8616 Qualification Tests 1.RoomTemperature Tensile a. 0.2% Y.S. 108 KSI UTS 126 KSI Elongation 6.0% b.0.2% Y.S. 112.2 KSI 111.0 KSI UTS 126 KSI 135.7 KSI Elongation 6.3% 6.7%2. Stress-Rupture Temperature Stress Rupture Life Elongation a. 1800 F.22 KSI 30.0 hours 1800 F. 24 KSI 14.8 hours 14.0% b. 1800 F. 22 KSI 55.3hours  9.1% 1800 F. 22 KSI 58.2 hours 10.3%

The room-temperature tensile and 1800 F stress-rupture properties of the28 percent recast INCO713C composite test specimens are summarized inTable III. TABLE III Measured Tensile and Stress-Rupture Properties ofComposite Cast/Recast INCO713C Test Specimens 1.Room Temperature TensileProperties Specimen 0.2 YS UTS Elongation #1 123.3 KSI 150.3 KSI 5.6% #2122.0 KSI 151.5 KSI 6.6% #3 122.4 KSI 148.1 KSI 6.7% Average 122.4 KSI150.0 KSI 6.3% Reduction in Area ® Specimen Rupture Life Elongation 1800F./22 KSI 2. Stress-Rupture Properties (stress calculated on castINCO713C cross-section only) #4 60.9 hrs. 10.7%  21.1% #5 55.9 hrs. 6.3%17.8% #6 60.9 hrs. 7.1% 16.8% @ 1600 F./42 KSI (stress calculated oncast INCO713C cross-section only) #5 202.5 hrs. 6.9% 12.2% #6 >212.5hrs. 4.9%  8.6%

The room temperature yield and ultimate tensile strengths of the 28percent recast INCO713C composite test specimens were approximately 11percent higher than those of as-cast INCO713C core material. The roomtemperature ductility of the 28 percent recast INCO713C composite testspecimens was virtually identical to that of the as-cast INCO713C corematerial.

The as-cast INCO713C core material and the 28 percent recast INCO713Ccomposite test specimens were tested for stress-rupture at 1800 F under“constant load” conditions to experimentally assess the effect of therecast process on the sustained, high-temperature, load-bearing capacityof as-cast INCO713C.

The approximate time to rupture as-cast INCO713C at 1800 F/22 KSI, asestimated from available “Larsen-Miller” correlations, is 48 hours. Thetime to rupture the as-cast INCO713C core material test bars at 1800F/22 KSI was 30.0 hours. The average time to rupture machined as-castINCO713C test specimens at 1800 F/22 KSI was 56.5 hours. The averageas-cast INCO713C 1800 F/22 KSI stress-rupture life was 45 hours, plus orminus 15 hours.

The 28 percent recast INCO713C composite test specimens were tested at1800 F under loads sufficient to produce 22 KSI stress based on as-castINCO713C substrate dimensions rather than composite test specimendimensions. Test loads ranged from 795 to 799 pounds (797 poundsaverage) depending on precise as-cast INCO713C machined diameters.Corresponding composite specimen stresses ranged from 15 to 16 KSI.

The average time to rupture the 28 percent INCO713C composite testspecimens under such “constant load” test conditions was 60.9 hours at1800 F.

Data Analyses:

The data summarized in Table III show that the recast process augmentsthe room temperature tensile properties of as-cast INCO713C.

Assuming the room temperature tensile properties of the as-cast INCO713Csubstrate remain unchanged by the thermal treatments associated with therecast process, “rule of mixture” analyses of the room temperature 28percent recast INCO713C composite tensile data summarized in Table IIIindicate that the recast INCO713C portion of the composite has thefollowing room temperature tensile properties: 150 KSI 0.2% yieldstrength 190 KSI ultimate tensile strength 5.8% elongation

The data summarized in Table III similarly show that the recast processaugments the sustained high-temperature, load-bearing capacity ofas-cast INCO7 1 3C.

“Load partitioning analysis”, for lack of a better description, wereused to distinguish the stress-rupture strength properties of the recastINCO713C coating from those of the as-cast INCO713C substrate.

“Larsen-Miller” stress-rupture data correlation's suggest that thestress required to increase the 1800 F rupture life of an as-castINCO713C substrate specimen to 60.9 hours is only 21 KSI. The loadrequired to develop a stress of 21 KSI, based on an average 0.2145 inchas-cast INCO713C substrate diameter, is 759 pounds. Since 797 poundswere applied to the 28 percent recast INCO713C composite specimenstested at 1800 F/16 KSI, it follows that the balance of the load (39pounds) was accommodated by the recast INCO713C coating.

Since the cross-sectional area of the recast INCO713C coating in the 28percent recast INCO713C composite specimens was 0.0161 square inches,the recast INCO713C coating stress was 2.4 KSI. The 1800 F/60.9 hourstress-rupture strength of recast INCO713C is, therefore, approximately2.4 KSI.

Two 28 percent recast INCO713C composite test specimens were similarlytested in stress-rupture at 1600 F under loads calculated to develop astress of 42 KSI based on as-cast INCO713C substrate dimensions.

One of the 28 percent recast INCO713C composite test specimens rupturedin 202.5 hours at 1600 F/42 KSI (based on as-cast substrate dimensions)while the other was arbitrarily terminated without rupture after 212.5hours. An as-cast INCO713C test specimen might be expected to rupture inapproximately 100 hours at 1600 F/42 KSI.

“Load-partitioning analyses” of these 1600 F stress-rupture test resultssuggest that the 1600 F/200 hour stress-rupture strength of the recastINCO713C coating is greater than 8 KSI.

The stress-rupture properties of the recast INCO713C coating, asinferred from “load partitioning analyses”, generally correspond tothose of wrought nickel or cobalt-base levels through post HIP heattreatments.

The experimental data discussed above indicate that recast INCO713Ccoating:

-   -   1. have intrinsically higher room temperature tensile strength        than as-cast INCO713C; and,    -   2. have intrinsic stress-rupture strengths approximately        equivalent to wrought nickel or cobalt-base alloys.

More importantly, the experimental data presented and discussed in thisstudy convincingly demonstrate that the recast process augments theroom-temperature tensile and sustained high-temperature, load-bearingcapacities of as-cast INCO713C.

In accordance with another aspect of the present invention, a method offorming metal products and components having a durable wear resistantcoating is provided. FIG. 1(b) is a flow chart showing the steps of theinventive method of forming metal products and metal components having awear resistant coating. This method obtains a metal product havingrobust diffusion bonding occurring between a metal substrate and anapplied coating. The first step of the inventive method is to determinethe attributes of a final workpiece product (Step One). For example, ifthe final workpiece product is a cutting tool the attributes include awear resistant surface formed on a relatively inexpensive tool substrate10. If the final workpiece is a cast metal component, a decorative,smooth final surface may be desired on a cast substrate 16.

An appropriate substrate composition is then determined (Step Two)depending on the selected attributes. In the example of a cutting tool,the substrate composition may be high speed steel, which is relativelyinexpensive to form but durable enough for its intended purpose. In thecase of a cast metal component, the cast workpiece substrate can beformed from cast iron or aluminum (or other cast metal or metal alloy).A workpiece substrate is formed to near-finished dimensions (StepThree), using known processes such as casting, extruding, molding,machining, etc. An appropriate coating material 12 composition isdetermined depending on the selected attributes (Step Four). Again, inthe example of a cutting tool the coating material 12 could be selectedfrom a number of relatively hard and durable metals and alloys such asCobalt, Carbide, TiN, etc. In the example of the cast metal component,aluminum oxide may be chosen to provide both a decorative and corrosionresistant surface. The selection of both the substrate and coatingcomposition also depends on their metallurgical compatibility with eachother.

The workpiece substrate is prepared for a high-density coating process(Step Five). The preparation may include cleaning, blasting, machining,masking or other like operations. Once the workpiece substrate has beenprepared, a high-density coating process is performed to coat theworkpiece substrate (Step Six). The coating material 12 is built-up to athickness that is effective to obtain desired finished dimensions afterperforming a hot isostatic pressing treatment (described below). Thehigh-density coating process may comprise performing a hyper velocityoxy-fuel thermal spray process. In the case of HVOF, a fuel gas andoxygen are used to create a combustion flame at 2500 to 3100° C. Thecombustion takes place at a very high chamber pressure and a supersonicgas stream forces the coating material 12 through a small-diameterbarrel at very high particle velocities. The HVOF process results inextremely dense, well-bonded coatings. Typically, HVOF coatings can beformed nearly 100% dense, with at a porosity of about 0.5%.

The high particle velocities obtained using the HVOF process results inrelatively better bonding between the coating material 12 and thesubstrate, as compared with other coating methods such as theConventional Plasma spray method or the Chemical Vapor Depositionmethod. However, the HVOF process also forms a bond between the coatingmaterial 12 and the substrate that occurs primarily through mechanicaladhesion at a bonding interface. As will be described below, inaccordance with the present invention this mechanical bond is convertedto a metallurgical bond by creating a diffusion bond between the coatingmaterial 12 and the workpiece substrate. The diffusion bond does nothave the interface boundary which is usually the site of failure.

The diffusion bond is created by subjecting the coated workpiecesubstrate to a hot isostatic pressing (HIP) treatment. The appropriatehot isostatic pressing treatment parameters are selected depending onthe coating, the workpiece substrate and the final attributes that aredesired (Step Seven). The hot isostatic pressing treatment is performedon the coated workpiece substrate to obtain a metal product having thedesired finished dimensions and diffusion bonding between the coatingmaterial 12 and the workpiece substrate (Step Eight).

By proper formation of the workpiece substrate, the final dimensions ofthe finished workpiece product can be accurately achieved through theprecise control of the build up of coating material 12 when the HVOFplasma spray process is performed. Alternatively, the HIP treated andHVOF coated workpiece substrate may be machined to final dimensions asnecessary (Step Nine).

HIP treatment is conventionally used in the densification of cast metalcomponents and as a diffusion bonding technique for consolidating powdermetals. In the HIP treatment process, a part to be treated is raised toa high temperature and isostatic pressure. Typically, the part is heatedto 0.6-0.8 times the melting point of the material comprising the part,and subjected to pressures on the order of 0.2 to 0.5 times the yieldstrength of the material. Pressurization is achieved by pumping an inertgas, such as Argon, into a pressure vessel 14. Within the pressurevessel 14 is a high temperature furnace, which heats the gas to thedesired temperature. The temperature and pressure is held for a setlength of time, and then the gas is cooled and vented.

The HIP treatment process is used to produce near-net shaped components,reducing or eliminating the need for subsequent machining operations.Further, by precise control of the temperature, pressure and time of aHIP treatment schedule a particular microstructure for the treated partcan be obtained.

In accordance with the present invention, the HIP treatment process isperformed on a HVOF coated substrate to convert the adhesion bond, whichis merely a relatively weaker mechanical bond, to a diffusion bond,which is a relatively stronger metallurgical bond. In accordance withthe present invention, an HVOF coating process is used to apply thecoating material 12 having sufficient density to effectively undergo thedensification changes that occur during the HIP process. A sinteringheat treatment step may be performed improve the density of the coatingmaterial and prevent gas entrapment during the hot isostatic pressingtreatment. If the coating material 12 and the workpiece substrate arecomprised of the same metal composition, then the diffusion bondingresults in a particularly seamless transition between the substrate andthe coating.

FIG. 1(c) is a flow chart showing the steps of the inventive method forcorrecting defects in a workpiece. A location of a defect in a workpieceis determined (Step one). The defect comprises, for example, a void oran inclusion in a workpiece substrate. For example, an oxide or dirtmight be introduced or formed in the workpiece during a manufacturingprocess. The workpiece substrate is comprised of a metal alloy. Materialof the workpiece substrate at the location of the defect is removed toform cleaned area in the workpiece substrate (Step two). The cleanedarea may be formed by sand or grit blasting, machining, grinding, or thelike. The cleaned area in the workpiece substrate is coated with ahigh-density coating (Step three). A sintering heat treatment isperformed on the coated workpiece substrate to remove entrapped gas fromthe coating material prior to a step of hot isostatic pressing treating(Step four). Then, hot isostatic pressing treating is performed on thecoated workpiece to produce diffusion bonding between the workpiecesubstrate and the high-density coating (Step five). If necessary, afterthe HIP process is complete, the coated workpeice may be machined to thedesired dimensions (Step six). A high-density coating process such ashyper-velocity oxy-fuel thermal spray process or a detonation gunprocess is used to apply the high-density coating to the substrate atthe location of the cleaned area. The high-density coating may have thesame metal alloy composition as the metal alloy substrate. The metalalloy substrate may comprise a nickel or cobalt-based superalloy, andthe high-density coating may have the same nickel or cobalt-based superalloy composition as the metal alloy substrate.

As shown in FIGS. 2(a) through 2(d), the inventive method can be usedfor forming a metal product having a wear resistant surface. FIG. 2(a)is a schematic view showing a tool substrate 10 provided in accordancewith the inventive method of forming metal components having a wearresistant coating. The inventive method can be employed to produce, forexample, a long lasting cutting tool from a relatively inexpensivecutting tool substrate 10.

In accordance with this aspect of the invention, a workpiece substrateis formed to near-finished dimensions. The tool substrate 10 may be adrill bit, end mill, lathe tool bit, saw blade, planer knifes, cuttingtool inserts, or other cutting tool part. The substrate may,alternatively, be something other than a tool. For example, ice skateblades and snow ski edges may be treated in accordance with the presentinvention to obtain a long wearing edge. Kitchen knives may be treatedin accordance with the present invention to reduce or even eliminate theneed for constant sharpening. Further, products such as pen tips andfishing hooks may be treated in accordance with the present invention soas to benefit from long lasting durability. Nearly any metal componentthat could benefit from a longer wearing, dense surface structure mightbe a candidate from the present invention. For example, steam turbineerosion shields, fly ash fan blades, power plant conveyors, are allsubjected to wear and/or surface erosion forces. The present inventioncan be used to provide the protective surface characteristics, asdescribed herein, that enhance the effectiveness of products such asthese.

FIG. 2(b) is a schematic view of the tool substrate 10 having a wearresistant coating applied using an HVOF thermal spray process inaccordance with the inventive method. A high-density coating process,such as a hyper velocity oxy-fuel thermal spray process, is performed tocoat the workpiece substrate 10 with a wear resistant coating material12 using, for example, an HVOF nozzle. The coating material 12 isbuilt-up to a thickness that is effective to obtain desired finisheddimensions after performing a hot isostatic pressing treatment.

FIG. 2(c) is a schematic view of the HVOF spray coated tool substrate 10undergoing a HIP treatment process in a HIP vessel 14. The hot isostaticpressing treatment is performed on the coated workpiece substrate toobtain a metal product having the desired finished dimensions anddiffusion bonding between the coating material 12 and the workpiecesubstrate.

FIG. 2(d) is a schematic view of the final HVOF spray coated and HIPtreated tool having a wear resistant coating layer diffusion bonded tothe tool substrate 10. In accordance with the present invention themechanical bond formed between the parent substrate and the appliedcoating is converted to a metallurgical bond by creating a diffusionbond between the coating material 12 and the parent substrate. Thediffusion bond does not have the interface boundary which is usually thesite of failure, thus a superior product is obtained that has desiredsurface properties, such as wear resistance, color, smoothness, texture,etc. These surface properties do not end abruptly at a bonding interface(as is the case of conventional coated or brazed products), but ratherremain present to a continuously varying degree from the product surfaceto the parent metal. A cutting edge can be put on the tool surface byconventional sharpening techniques taking care not to remove more of thediffusion bonded coating than is necessary.

FIGS. 3(a) through 3(e) illustrate the present inventive method employedfor forming a cast metal product having predetermined dimensions andsurface characteristics. FIG. 3(a) is a schematic perspective view of acast metal workpiece substrate undergoing a machining operation. Asshown in FIG. 3(a), the cast metal workpiece is machined, if necessary,to near-finished dimensions. FIG. 3(b) is a schematic perspective viewof the machined cast metal component.

A high-density coating process, such as a hyper velocity oxy-fuelthermal spray process, is performed to coat the workpiece substrate witha coating material 12. FIG. 3(c) is a schematic perspective view of themachined cast metal component having a coating applied using an HVOFthermal spray process. The coating material 12 is built-up to athickness effective to obtain desired finished dimensions afterperforming a hot isostatic pressing treatment. FIG. 3(d) is a schematicperspective view of the HVOF spray coated machined cast metal componentundergoing a HIP treatment process in a HIP vessel 14. The hot isostaticpressing treatment is performed on the coated workpiece substrate toobtain a metal product having the desired finished dimensions anddiffusion bonding between the coating material 12 and the workpiecesubstrate. FIG. 3(e) is a schematic perspective view of the final HVOFspray coated and HIP treated machined cast metal product having acoating layer diffusion bonded to the machined cast metal component.

FIG. 4 is a flow chart showing the steps of the inventive method ofrepairing a turbine engine part. The present inventive method can beused for repairing a turbine engine part 18, such as a blade or vane. Inaccordance with this aspect of the invention a turbine engine part 18,which is comprised of a metal or metal alloy, is first cleaned (StepOne). If necessary, eroded portions of the turbine engine part 18 arewelded using a weld material comprised of the same metal or metal alloyas the parent or original metal engine part (Step Two). The weldingoperation is performed to build up heavily damaged or eroded portions ofthe turbine engine part 18. If the part is not heavily damaged, thewelding operation may be obviated.

The welding operation will typically produce weld witness lines. Theweld witness lines are ground flush to prevent blast material frombecoming entrapped in the weld witness lines (Step Three). Portions ofthe engine part that are not to be HVOF sprayed are masked (Step Four),and the engine part is again cleaned in preparation for HVOF spraying(Step Five). HVOF plasma spraying of the unmasked portions of the enginepart is performed (Step Six). The HVOF plasma spray material (coatingmaterial 12) is comprised of the same metal alloy as the parent ororiginal metal engine part. The HVOF plasma spray material is applied soas to build up a cordal dimension of the engine part to a thicknessgreater than the thickness of an original cordal dimension of the enginepart. A sintering heat treatment process may be performed to furtherdensify the coating material. A hot isostatic pressing (HIP) treatmentif performed on the coated engine part to densify the coating material12, to create a diffusion bond between the coating material 12 and theparent and weld material, and to eliminate voids between the turbineengine part 18, the weld material and the coated material (Step Seven).Finally, the engine part is machined, ground and/or polished to theoriginal cordal dimension (Step Eight).

FIG. 5(a) is a schematic side view and FIG. 5(b) is a schematiccross-sectional view of a worn turbine engine part 18 before undergoingthe inventive method of repairing a turbine engine part 18. Metal alloycomponents, such as gas turbine parts such as blades and vanes, areoften damaged during use. During operation, gas turbine parts aresubjected to considerable degradation from high pressure and, in thecase of rotating components such as blades, centrifugal force in a hotcorrosive atmosphere. The gas turbine parts also sustain considerabledamage due to impacts from foreign particles. Further, during inspectionand/or repair operations the engine parts are stripped of a protectivediffusion coating, which usually results in the reduction of some of thesubstrate thickness. This degradation results in a limited service lifefor these parts. Since they are costly to produce, various conventionalrepair methods are employed to refurbish damaged gas turbine blades andvanes. However, these conventional repair methods generally requirelabor intensive machining and welding operations that often subject thepart to damaging stress. Also, these conventional repair methodstypically utilize low pressure plasma spray for the application of acoating material 12. Conventional plasma spray coating methods depositthe coating material 12 at relatively low velocity, resulting in voidsbeing formed within the coating and in a coating density typicallyhaving a porosity of about 5.0%. Again, the bond between the substrateand the coating occurs primarily through mechanical adhesion at abonding interface, and if the coating is subjected to sufficientshearing forces it will flake off of the workpiece substrate. Further,the high porosity of the coating obtained through conventional plasmaspray coating make them inadequate candidates for diffusion bondingthrough the HIP treating process described herein.

FIG. 6(a) is a schematic side view and FIG. 6(b) is a schematiccross-sectional view of the worn turbine engine part 18 showing the wornareas 20 to be repaired using the inventive method of repairing aturbine engine part 18. The area enclosed by the dashed lines representthe material that has been erode or otherwise lost from the originalturbine engine part 18. In accordance with the present invention, thisarea is reconstituted using the same material as the original blade andusing the inventive metal treatment process. The worn turbine enginepart 18 (in this case, a turbine blade) is first cleaned to prepare theworn surfaces for welding (see Step One, FIG. 4).

FIG. 7(a) is a schematic side view and FIG. 7(b) is a schematiccross-sectional view of the worn turbine engine part 18 showing the wornareas filled in with similar weld material 22 in accordance with theinventive method of repairing a turbine engine part 18 (see Step Two,FIG. 4). In accordance with the present invention, the weld material isthe same as the original blade material making the bond between the weldand the substrate exceptionally strong.

FIG. 8(a) is a schematic side view and FIG. 8(b) is a schematiccross-sectional view of the welded turbine engine part 25 showing areas24 to be built up with similar coating material 12 using an HVOF spraycoating process in accordance with the inventive method of repairing aturbine engine part. In accordance with the present invention, thecoating material 12 is the same as the original blade material, againmaking the bond between the weld and the substrate exceptionally strong.

FIG. 9(a) is a schematic side view and FIG. 9(b) is a schematiccross-sectional view of the HVOF built up, welded turbine engine part 27showing an area, such as the vane or blade root, masked 26 beforeperforming the HVOF spray coating process in accordance with theinventive method of repairing a turbine engine part. The coatingmaterial 12 is built-up to a thickness that is effective to obtaindesired finished dimensions after performing a hot isostatic pressingtreatment (described below).

The high-density coating process may comprise performing a hypervelocity oxy-fuel thermal spray process. In the case of HVOF, a fuel gasand oxygen are used to create a combustion flame at 2500 to 3100° C. Thecombustion takes place at a very high chamber pressure and a supersonicgas stream forces the coating material 12 through a small-diameterbarrel at very high particle velocities. The HVOF process results inextremely dense, well-bonded coatings. Typically, HVOF coatings can beformed nearly 100% dense, at a porosity of about 0.5%. The high particlevelocities obtained using the HVOF process results in relatively betterbonding between the coating material 12 and the substrate, as comparedwith other coating methods such as the conventional plasma spray methodor the chemical vapor deposition method. However, the HVOF process formsthe bond between the coating material 12 and the substrate that occursprimarily through mechanical adhesion at a bonding interface. As will bedescribed below, in accordance with the present invention thismechanical bond is converted to a metallurgical bond by creating adiffusion bond between the coating material 12 and the workpiecesubstrate. The diffusion bond does not have the interface boundary whichis usually the site of failure.

The diffusion bond is created by subjecting the coated workpiecesubstrate to a hot isostatic pressing (HIP) treatment. The appropriatehot isostatic pressing treatment parameters are selected depending onthe coating, the workpiece substrate and the final attributes that aredesired. The hot isostatic pressing treatment is performed on the coatedworkpiece substrate to obtain a metal product having the desiredfinished dimensions and diffusion bonding between the coating material12 and the workpiece substrate.

FIG. 10 is a schematic view of the HVOF built up, welded turbine enginepart 27 undergoing a HIP treatment process in a HIP vessel 14 inaccordance with the inventive method of repairing a turbine engine part.

HIP treatment is conventionally used in the densification of cast metalcomponents and as a diffusion bonding technique for consolidating powdermetals. In the HIP treatment process, a part to be treated is raised toa high temperature and isostatic pressure. Typically, the part is heatedto 0.6-0.8 times the melting point of the material comprising the part,and subjected to pressures on the order of 0.2 to 0.5 times the yieldstrength of the material. Pressurization is achieved by pumping an inertgas, such as Argon, into a pressure vessel 14. Within the pressurevessel 14 is a high temperature furnace, which heats the gas to thedesired temperature. The temperature and pressure is held for a setlength of time, and then the gas is cooled and vented.

The HIP treatment process is used to produce near-net shaped components,reducing or eliminating the need for subsequent machining operations.Further, by precise control of the temperature, pressure and time of aHIP treatment schedule a particular microstructure for the treated partcan be obtained.

FIG. 11(a) is a schematic side view and FIG. 11(b) is a schematiccross-sectional view of the final HVOF spray coated and HIP repairedturbine engine part 28 having a similar metal coating layer diffusionbonded to the original parent substrate and welded portions inaccordance with the inventive method of repairing a turbine engine part.By proper formation of the workpiece substrate, the final dimensions ofthe finished workpiece produce can be accurately achieved through theprecise control of the build up of coating material 12 when the HVOFplasma spray process is performed. Alternatively, the HIP treated andHVOF coated workpiece substrate may be machined to final dimensions asnecessary (Step Eight).

An experimental test piece was prepared in accordance with the inventivemethod of treating metal components. Photomicrographs of the test pieceshowed the grain structure and diffusion bonding of the coating material12 and the substrate after the inventive method has been performed. TheHIP treatment process was performed on an HVOF coated test substrate toconvert the adhesion bond between the coating and the substrate, whichis merely a mechanical bond, to a diffusion bond, which is ametallurgical bond. In accordance with the present invention, an HVOFcoating process is used to apply the coating material 12 havingsufficient density to effectively undergo the densification changes thatoccur during the HIP process. In the case of the test piece example, thecoating material 12 and the workpiece substrate are comprised of thesame metal composition. The diffusion bonding results in a transitionbetween the substrate and the coating that has a much strongerstructural integrity and wear characteristics as compared with theconventional art.

The test piece was prepared by building up coating material 12 to athickness of approximately 0.02 inches, and the composition of the testpieces was determined at seven locations (A-G) across a cross section ofthe piece. The composition was found to be substantially uniform acrossthe cross-section of the test piece, as shown in the following table.TABLE I Elemental Composition (Weight %) Element A B C D E F G Aluminum5.4 5.2 5.5 6.2 6.3 6.4 6.5 Titanium 0.6 0.6 1.0 0.6 1.0 0.6 0.9Chromium 12.9 13.2 14.5 12.7 11.5 13.7 14.1 Nickel REM REM REM REM REMREM REM Niobium 1.4 1.5 1.8 2.1 1.7 2.3 2.6 Molybdenum 3.7 4.1 3.6 3.33.4 3.9 3.0

A photomicrograph of the treated workpiece shows the grain structure anddiffusion bonding of the coating material 12 and the substrate after theinventive method has been performed. In accordance with the presentinvention, the HIP treatment process is performed on a HVOF built up,welded turbine engine part to convert the adhesion bond, which is merelya mechanical bond, to a diffusion bond, which is a metallurgical bond.In accordance with the present invention, an HVOF coating process isused to apply the coating material 12 having sufficient density toeffectively undergo the densification changes that occur during the HIPprocess. If the coating material 12 and the workpiece substrate arecomprised of the same metal composition, then the diffusion bondingresults in smooth transition between the substrate and the coating. Incontrast, a conventional plasma spray coating method results in arelatively weak bond between the coating and the substrate. The bond isprimarily due to a mechanical adhesion bond that occurs relativelylocally within a boundary interface.

As discussed in detail above, in accordance with the present inventivemethod a deformed gas turbine engine airfoil part can be returned to thedimensions required to place the part back into useful service. Adiffusion bond is created between the coating material and the substrateof a repaired gas turbine engine airfoil part. This diffusion bond isextremely robust and results in a repaired engine part that has theappropriate mechanical properties that allow the part to be safelyreturned to service. The inventive method of repairing a turbine engineairfoil part offers substantial savings because it provides for theefficient and effective repairing of expensive engine parts whichotherwise might have been discarded.

As shown in FIG. 13 in accordance with another aspect of the presentinvention, the reclassification of a gas turbine engine airfoil part isobtained. The dimensional differences between pre-reclassifieddimensions of the buttresses of a turbine engine airfoil part anddesired post-reclassified dimensions of the buttresses are determined(Step One). That is, the change in shape of the inner buttress and outerbuttress necessary to obtained a desired angular relationship betweenthe airfoil section and the buttresses is determined. Build-up thicknessof coating material required to obtain the desired post-reclassifieddimensions of the buttresses is determined (Step Two). A high-densitycoating process, such as HVOF, is used to coat the buttresses of theturbine engine airfoil part with a coating material (Step Three). Theportions of the part that are not to be built up, such as the airfoilsection and parts of the buttresses, may be masked before applying thehigh-density coating. Also, some of the coated surfaces of the part mayneed to be built up more than others. The coating material is applied atleast to the determined build-up thickness of coating material effectiveto obtain the desired post-reclassification dimensions after performinga hot isostatic pressing treatment, and after the selective removal ofsome of the original buttress material and some of the built up coatingmaterial.

As discussed herein, the coating material comprises a metal alloycapable of forming a diffusion bond with the substrate of the turbineengine airfoil part. After the coating material is applied, thesintering heat treatment process may be performed (Step Four) to preventgas entrapment of the coating material and/or the diffusion bonding areaduring the hot isostatic pressing process. Then, the hot isostaticpressing process is performed so that the buttresses of the turbineengine airfoil part have a robust diffusion bonding between the coatingmaterial and the original material of the buttresses (Step Five). Havingbuilt up the appropriate dimensions of the inner buttress and outerbuttress, the reclassification of the part is obtained by selectivelyremoving the original buttress material and, if necessary, some of thebuilt up material until the angular relationship between the airfoilsection and the inner and outer buttresses is obtained (Step Six). Thematerial can be removed through milling, grinding, or other suitable andwell known machining operations. Further, to facilitate obtaining thecorrect dimensions the centerline position of the airfoil part can belocated and held by mounting the part in a suitable holding fixture whenmachining the buttresses.

The fixture may be so constructed so that a vane that has at least aminimum amount of material built up on its buttresses can be machinedand reclassified. In this case, it may not be necessary to determine thedimensional differences or the required build-up thickness. Rather, theinventive high density coating and HIPing process (and, if neededsintering) can be performed to build up at least the minimum amount ofmaterial diffusion bonded to the buttresses. Then, the vane is placed inthe fixture and the excess material (both original buttress material andthe built-up material) is machined until the buttresses have beenreshaped and the vane reclassified as intended or restored to original.

The class of a turbine engine vane is defined by the angularrelationship between the airfoil section and the inner and outerbuttresses. The inventive recast process is utilized to change orrestore the original class of a turbine engine airfoil part by buildingup sufficient material on the inner buttress and the outer buttress sothat the buttresses can then be machined to create the desired angles αand α′ (shown in FIGS. 14(b) and 14(c)) and reclassify the vane.

All buttresses are dimensionally the same and all airfoils aredimensionally the same for all classes of vanes. In accordance with thepresent invention, the airfoil centerline position is held by mountingthe vane in a fixture, and the buttresses are machined to obtained todesired reclassification parameters.

The class of a turbine engine vane 20 is defined by the angularrelationship between the airfoil section 22 and the inner buttress 24and outer buttress 26. The inventive recast process is utilized tochange or restore the original class of a turbine engine airfoil part bybuilding up sufficient material on the inner buttress 24 and the outerbuttress 26 so that the buttresses 24, 26 can then be machined to createthe desired angles α and α′ (shown in FIGS. 14(b) and 14(c)) andreclassify the vane 20.

All buttresses 24, 26 are dimensionally the same and all airfoils aredimensionally the same for all classes of vanes. In accordance with thepresent invention, the airfoil centerline position is held by mountingthe vane 20 in a fixture, and the buttresses 24, 26 are machined toobtained to desired reclassification parameters.

FIG. 14(a) is a front view of a vane 20 from a gas turbine engineshowing the airfoil section 22, the outer buttress 26 and the innerbuttress 24. In accordance with this aspect of the invention, it isfirst determined what dimensions of the inner buttress 24 and outerbuttress 26 need to be adjusted in order to obtain the desiredreclassification of the vane 20. Having determined the dimensionaldifferences between the pre-reclassified buttresses 24, 26 and thepost-reclassified buttresses 24, 26, it is next to determine how muchmaterial must be added, and where the material must be added so that thebuttresses 24, 26 can be reshaped.

FIG. 14(b) is a partial top view showing the outer buttress 26 and angleα indicating the angular relationship between the airfoil section 22 andthe outer buttress 26 and FIG. 14(c) is a partial bottom view showingthe inner buttress 24 and angle α′ indicating the angular relationshipbetween the airfoil section 22 and the inner buttress 24. In accordancewith the present invention, the vane 20 is reclassified by changing theshape of the buttresses 24, 26 so that the angles α and α′ are changedresulting in a changed angle of attack of the airfoil section 22, andthus reclassification of the vane 20.

FIG. 14(d) is a partial left-side view showing the leading edge foot 28of the inner buttress 24 and the outer foot front face 30 of a buttressrail 32 of the outer buttress 26 and FIG. 14(e) is a partial right-sideview showing the trailing edge foot 34 of the inner buttress 24 and theother buttress rail 32 of the outer buttress 26. In accordance with thepresent invention, the surfaces of the buttresses 24, 26, such as theleading edge foot 28, center log 36, trailing edge foot 34 (innerbuttress 24), and the outer foot front face 30 and buttress rails 32(outer buttress 26) are selectively built up and machined so that theangle of attack of the airfoil section 22 is adjusted. The build up ofmaterial on the buttresses 24, 26 may be uniform, and then thebuttresses 24, 26 machined to selectively remove portions of theoriginal substrate and portions of the build up material. To reducemachine costs, the surfaces of the original buttresses 24, 26 that aregoing to be machined may be masked before the buildup material isapplied. In this case, the buildup material will not have to be latermachined along with the original substrate to reshape the buttresses 24,26, 24, 26.

A fixture for holding the vane 20 during the machining operation(s) maybe so constructed so that the vane 20 having at least a minimum amountof material built up on its buttresses 24, 26 can be machined andreclassified. In this case, it may not be necessary to determine thedimensional differences or the required build-up thickness. Rather, theinventive high density coating and HIPing process (and, if neededsintering and other processes described herein) can be performed tobuild up at least the minimum amount of material diffusion bonded to thebuttresses 24, 26, 24, 26. Then, the vane 20 is placed in the fixtureand the excess material (both original buttress material and thebuilt-up material) is machined until the buttresses 24, 26 have beenreshaped and the vane reclassified as intended.

The resulting reclassified vane has inner and outer buttresses with themechanical properties required for safe return to active service in anoperating gas turbine engine. The diffusion bonding between the appliedcoating material built up on the buttresses and the original buttresssubstrate ensures, as substantiated by the test results discussedherein, that the reclassified vane can be safely returned to activeservice.

FIG. 15(a) is a flowchart showing the steps of the inventive method forrepairing a workpiece with an electroplated coating diffusion bonded tothe workpiece. A workpiece substrate is provided and prepared for acoating operation (Step One). The preparation may include, for example,masking off portions that are not to be coated, cleaning and machiningsurfaces to be coated, etc. A coating is formed on at least selectedportions of the workpiece substrate through an electroplating process(Step Two). The coating material is capable of forming a diffusion bondwith the workpiece substrate. The diffusion bond is a metallurgical bondbetween the workpiece and the coating that does not have an interfaceboundary. This diffusion bond creates a secure attachment between thecoating and the substrate, much stronger than the mechanical bond thatis originally formed between the coating and the substrate. Thisdiffusion bond is formed through the hot isostatic pressing treatment.The diffusion bond can be formed when the coating on the substrate isdense. It may be possible to form this coating by a spray process, suchas vacuum spray, detonation gun, HVOF, or by a solution process such aselectroplating. To ensure a diffusion bond is formed, a sintering heattreatment may have to be first performed to densify the coating prior tothe hot isostatic heat treament step and, if necessary, to removeentrapped gas (Step Three). If the coating is not dense enough, it mayflake off of the substrate during the heat and pressure of the hotisostatic treatment step. Further, applicant has found that entrappedgas is problematic because it results in a weaker, bubbled surface withan inconsistent diffusion bond between the coating and the substrate.The sintering heat treatment densifies the coating and removes entrappedgas and prevents outgassing of the trapped gas during a hot isostaticpressing treatment. This preventive treatment has been experimentallyproven to greatly reduces the formation of bubbles on the surface of thecoated workpiece after the hot isostatic pressing treatment. After theentrapped gas is removed by the sintering heat treatment, the hotisostatic pressing treatment is then performed to drive the coatingmaterial into the workpiece substrate (Step Four). The hot isostaticpressing treatment results in the formation of the diffusion bond sothat the metallurgical bond between the workpiece and the coating isformed. Further post-HIP treatments can be performed such as heattreatments, machining operations, removing masking material, forming aprotective coating over the diffusion bonded coating, etc (Step Five).

FIG. 15(b) is a flow chart showing the steps of the inventive method forrepairing a gas turbine engine airfoil part with an electroplatedcoating diffusion bonded to the airfoil substrate. In accordance withthe present invention, the protective coating on a turbine engineairfoil part is removed so that the part can be prepared for theinventive electroplating recast repair method (Step One). Thedimensional differences between pre-repaired dimensions of a turbineengine airfoil part and desired post-repair dimensions of the turbineengine airfoil part are determined (Step Two). The turbine engineairfoil part has a substrate comprised of a superalloy. A build-upthickness of coating material required to obtain the desired post-repairdimensions of the turbine engine airfoil part is determined (StepThree). An electroplating process is used to coat the turbine engineairfoil part with a coating material to the determined build-upthickness of coating material effective to obtain the desiredpost-repair dimensions after performing a sintering heat treatment and ahot isostatic pressing treatment (Step Four). The electroplating processallows the controlled build up of material even between surfaces andaround angles of the substrate that would be difficult or impossible tocoat using a spray coating process. A spray coating process requires astraight line from the spary nozzle to the coated surface. When thesurface has contours and/or interior portions it is difficult orimpossible to coat these surfaces using a spray coating process. Even ifthe coating material can be sprayed into the contour or interiorportion, it remains difficult or impossible to apply an even coatingthickness. The electroplating process enables the coating to be appliedevenly even within interior surfaces, around corners or onto contours.The coating material comprises a metal alloy capable of forming adiffusion bond with the substrate of the turbine engine airfoil part.After the coating material is applied, a sintering heat treatmentprocess may be performed if necessary to densify the electroplatedcoating prior to the hot isostatic pressing process (Step Five). Theelectroplating process has the advantages of enabling a uniform coatingto be applied to a substrate, even if the substrate has contours andinterior spaces. The electroplating process may not result in trappedgas, as a spray coating process does. However, it still may be necessaryto perform the sintering heat treatment in order to densify the coating,so as to prevent the coating from flaking from the substrate due to theheat and pressure of the hot isostatic pressing treatment. The hotisostatic pressing process is performed to obtain a post-repair turbineengine airfoil part having the desired post-repair dimensions and havingdiffusion bonding between the coating material and the turbine engineairfoil substrate (Step Six). After performing the hot isostaticpressing process, a protective coating may be re-applied (Step Seven).Typically, this protective coating is present on an airfoil part toprotect it from the hot corrosive environment it experiences duringservice. This protective coating must be removed during the inspectionand/or repair process. After undergoing a number of inspection and/orrepair cycles, the airfoil part was conventionally discarded simplybecause the airfoil dimensions of the part were too deformed for thepart to be usable. However, in accordance with the present inventiverepair method, the airfoil dimensions are restored and a robust repairedairfoil part is obtained

In the typical application of the inventive method, the metal alloysubstrate of the turbine engine airfoil part will comprise a nickel orcobalt-base superalloy. The step of performing the electroplatingcoating process (Step Four) may include performing the electroplatingcoating process using an electroplatable material that is effective tocreate a diffusion bond with the airfoil substrate after the sinteringand hot isostatic pressing treatment steps.

By performing the inventive method for repairing a gas turbine engineairfoil part, the post-repair dimensions are equal to the dimensionsnecessary for effectively returning the part to active service. Theobtained diffusion bonding between the coating material and thesubstrate ensures that the repaired airfoil part is robust enough towithstand the highly demanding environmental conditions present in anoperating gas turbine engine. Thus, the present invention offerssubstantial cost savings over having to replace a turbine gas engineairfoil part which otherwise might have been discarded. The presentinvention can be used as a process for restoring critical gas path areadimensions in cast nickel or cobalt-base superalloy vane components.These dimensions may become altered due to erosion or particle strikesduring the service life of the part, and/or may become altered during aninspection or repair process wherein a protective coating is strippedfrom the part.

FIG. 15(c) is a flow chart showing the steps of the inventive method forcorrecting defects in a workpiece with an electroplated coatingdiffusion bonded to the workpiece. A location of a defect in a workpieceis determined (Step one). The defect may comprise, for example, a voidor an inclusion in a workpiece substrate. For example, a crack or divotmay be present in the workpiece due to manufacturing or service-relatedproblems. Or, an oxide or dirt might be introduced or formed in theworkpiece during a manufacturing process. Further, a cast workpiece mayhave casting flaws such as surface porosity, voids, cracks, or may beundersized due to shrinkage. The invention method for correcting defectsin a workpiece can be employed to correct such casting defects prior tofinish machining operations. The workpiece substrate is comprised of ametal alloy. Material of the workpiece substrate at the location of thedefect may be removed, if necessary, to form a cleaned area in theworkpiece substrate (Step two). The cleaned area may be formed by sandor grit blasting, machining, grinding, selective etching, or the like.Parts of the workpiece that are not to be coated may then be masked.

An electroplating process is used to coat the turbine engine airfoilpart with a coating material to the determined build-up thickness ofcoating material effective to obtain the desired post-repair dimensionsafter performing a sintering heat treatment and a hot isostatic pressingtreatment. The electroplating process allows the controlled build up ofmaterial even between surfaces and around angles of the substrate thatwould be difficult to coating using a spray process. The coatingmaterial comprises a metal alloy capable of forming a diffusion bondwith the substrate of the turbine engine airfoil part (Step three). Asintering heat treatment may be performed on the coated workpiecesubstrate to densify the coating material prior to a step of hotisostatic pressing treatment (Step four). Then, hot isostatic pressingtreatment is performed on the coated workpiece to produce diffusionbonding between the workpiece substrate and the electroplated coating(Step five). If necessary, after the HIP process is complete, themasking may be removed and/or the coated workpiece may be machined tothe desired dimensions (Step six).

FIG. 15(d) is a flow chart showing the steps of the inventive method forreclassifying a gas turbine engine airfoil part with an electroplatedcoating diffusion bonded to the airfoil part. The dimensionaldifferences between pre-reclassified dimensions of the buttresses of aturbine engine airfoil part and desired post-reclassified dimensions ofthe buttresses are determined (Step One). That is, the change in shapeof the inner buttress and outer buttress necessary to obtained a desiredangular relationship between the airfoil section and the buttresses isdetermined. Build-up thickness of coating material required to obtainthe desired post-reclassified dimensions of the buttresses is determined(Step Two).

An electroplating process is used to coat the turbine engine airfoilpart with a coating material to the determined build-up thickness ofcoating material effective to obtain the desired post-repair dimensionsafter performing a sintering heat treatment and a hot isostatic pressingtreatment. The electroplating process allows the controlled build up ofmaterial even between surfaces and around angles of the substrate thatwould be difficult to coat using a spray process. The coating materialcomprises a metal alloy capable of forming a diffusion bond with thesubstrate of the turbine engine airfoil part (Step three). The portionsof the part that are not to be built up, such as the airfoil section andparts of the buttresses, may be masked before applying the electroplatedcoating. Also, some of the coated surfaces of the part may need to bebuilt up more than others. In this case, the masking can be done instages, so that after a build up of electroplated material occurs, aportion of the built up surface is masked before additionalelectroplating build is performed on the unmasked portions. The coatingmaterial is applied at least to the determined build-up thickness ofcoating material effective to obtain the desired post-reclassificationdimensions after performing a hot isostatic pressing treatment, andafter the selective removal of some of the original buttress materialand some of the built up coating material.

As discussed herein, the coating material comprises a metal alloycapable of forming a diffusion bond with the substrate of the turbineengine airfoil part. After the coating material is applied, thesintering heat treatment process may be performed (Step Four) to densifythe electroplated coating prior to the hot isostatic pressing process.Then, the hot isostatic pressing process is performed so that thebuttresses of the turbine engine airfoil part have a robust diffusionbonding between the coating material and the original material of thebuttresses (Step Five). Having built up the appropriate dimensions ofthe inner buttress and outer buttress, the reclassification of the partis obtained by selectively removing the original buttress material and,if necessary, some of the built up material until the angularrelationship between the airfoil section and the inner and outerbuttresses is obtained (Step Six). The material can be removed throughmilling, grinding, or other suitable and well known machiningoperations. Further, to facilitate obtaining the correct dimensions thecenterline position of the airfoil part can be located and held bymounting the part in a suitable holding fixture when machining thebuttresses.

FIG. 16(a) shows an airfoil part prepared for electroplating. Aworkpiece substrate is provided and prepared for a coating operation.The preparation may include, for example, masking-off portions that arenot to be coated, cleaning and machining surfaces to be coated, etc

FIG. 16(b) shows the prepared airfoil part being electroplated. Acoating is formed on at least selected portions of the workpiecesubstrate through an electroplating process. The coating material iscapable of forming a diffusion bond with the workpiece substrate. Thediffusion bond is a metallurgical bond between the workpiece and thecoating that does not have an interface boundary. This diffusion bondcreates a secure attachment between the coating and the substrate, muchstronger than the mechanical bond that is originally formed between thecoating and the substrate.

FIG. 16(c) shows the electroplated airfoil part undergoing a sinteringheat treatment. A sintering heat treatment may be performed to densifythe coating material (Step Three). The sintering heat treatment may benecessary to prevent the coating material from separating from theworkpiece substrate under the temperature and pressure of the hotisostatic heat treatment.

FIG. 16(d) shows the sintered electroplated airfoil part undergoing ahot isostatic heat treatment. After the sintering heat treatment, thehot isostatic pressing treatment is then performed to drive the coatingmaterial into the workpiece substrate (Step Four). The hot isostaticpressing treatment results in the formation of the diffusion bond sothat the metallurgical bond between the workpiece and the coating isformed.

FIG. 16(e) shows the finished airfoil part having a diffusion bondbetween the electroplated areas and the airfoil substrate. Furtherpost-HIP treatments can be performed such as heat treatments, machiningoperations, removing masking material, forming a protective coating overthe diffusion bonded coating, etc (Step Five).

FIG. 17 illustrates the steps of correcting the dimensionalcharacteristics of a cast article. In accordance with the presentinvention, a method of correcting the dimensional characteristics of acast article is provided. As shown in Step One, the dimensionaldifferences are determined between pre-repair cast article dimensionsand desired post repair cast article dimensions to correct a castingdefect in the article. The determination may be made by determining thelocation of a void(s) in the surface of the article. The determinationmay also be made by determining an amount of buildup volume required tomake at least a portion of the surface of the cast article built up tothe desired post repair dimensions. For example, as shown in step one, adefect consisting of an inclusion can be found on the surface of a castarticle. As shown in step two, the inclusion is removed, and thesubstrate material in the immediate area around where the inclusion washas been removed by a machining operation such as drilling or milling.As shown in step three, the article is coated in at least an area of thecasting defect with a high-density coating material capable of forming adiffusion boundary between the coating material and the article. Asintering heat treatment may be performed to remove any trapped gasand/or to densify the coating surface to prevent gas from infiltratingthe coating during a hot isostatic pressing treatment (step four). Asshown in step five, the hot isostatic heat treatment process isperformed to form the diffusion boundary between the coating materialand the article. By this method, casting defects, such as oxideinclusions, surface bubbles or undercastings can be repaired. Therepaired area has filler material diffusion bonded with the castingsubstrate, ensuring the integrity of the repair.

Depending on the type of casting defect, material in an area of thecasting defect may be removed before the step of coating the article.For example, if the casting defect is an inclusion of an undesiredcomposition, such as an oxide or dirt particle, the inclusion and someof the base article material can be removed by a machining or otheroperation (step two). The area of the casting defect is enlarged, andmay be contoured to create a better surface for holding the coatingmaterial. The casting defect may be caused, for example, by at least oneof an inclusion at a surface of the article, an air bubble at thesurface of the article, undercasting, a void and shrinkage. A sinteringheat treatment can be performed (step four) before the step ofperforming the hot isostatic heat treatment (step five) to limit theoccurrence bubbles on the surface of the coating material after anisostatic heat treatment. The sintering heat treatment may preferrablybe performed at a temperature substantially the same as the temperatureof the hot isostatic heat treatment.

In accordance with the present invention, the coating material maycomprise an alloy with substantially no oxide forming constituents so asto avoid the formation of oxide inclusions in the coating material. Inthis case, the coating material may be applied using a coating processthat is effective to create a coating on the surface of the article thatwill be diffusion bonded to the article after the hot isostatic heattreatment, without the formation of crack inducing oxides. Applicant hasdiscovered that by preventing the formation of oxide constituents in thecoating, the ductility and other desirable properties of the coating isimproved. This improved ductility provides a protective barrier that mayeffectively prevent the propagation and the formation of cracks in thecoating material, the diffusion boundary and the substrate. An exampleof the chemistry of a suitable non-oxide inclusion forming coating is asfollows: Element Percentage Nickel Balance Chromium 9.0 Cobalt 10.0Carbon 0.14 Molybdenum 8.6 Tungsten 12.5 Boron 0.015 Columbium 1.0

As shown in step six, the coated surface of the substrate may besmoothed using a grinding or polishing operation. Thus, in accordancewith the present invention, the dimensional characteristics of the castarticle are corrected.

FIG. 18 is a flow chart showing the steps of the inventive method ofcorrecting the dimensional characteristics of a cast article and forproviding a protective coating to a metal article. A defect (which maybe a casting defect or other defect such as wear and tear) is identified(step one). Material may be removed from the workpiece substrate asnecessary (step two). For example, correcting a defect may require thatthe inclusion material and substrate material surrounding the inclusionbe removed. A bubble may leave a semi-spherical pit which can be drilledfor easier filling with the coating material. A coating material isapplied at least to the area of the defect. The coating material iscapable of forming a diffusion boundary between the coating material andthe article. In accordance with this aspect of the invention, thecoating material comprises an alloy with substantially no oxide formingconstituents so as to avoid the formation of oxide inclusions in thecoating material (step three). Applicants have discovered that theoxides in the coating may form crack initiation sites, and cracks formeddue to the oxides may propagate through the diffusion boundary and intothe article substrate. By limiting the formation of oxides in thecoating, these crack initiation sites are reduced or eliminated, therebyenabling the coating material to act as a protective coating. Asintering heat treatment can be performed before the step of performingthe hot isostatic heat treatment to limit the occurrence bubbles on thesurface of the coating material after an isostatic heat treatment. Thesintering heat treatment may be performed at a temperature substantiallythe same as the temperature of the hot isostatic heat treatment (stepfour). The hot isostatic heat treatment process is performed to form thediffusion boundary between the coating material and the article (stepfive). The coated surface of the substrate may be smoothed using agrinding or polishing operation (step six). Thus, in accordance with thepresent invention, the dimensional characteristics of the cast articleare corrected, as necessary, and the substantially oxide free coatingand the diffusion boundary provide a protective coating to protect thearticle from damage. FIG. 19 schematically illustrates a coatedsubstrate wherein the coating

material is diffusion bonded to the substrate and includes an oxideinclusion. FIG. 20 schematically illustrates the coated substrate shownin FIG. 19 wherein a crack is forming at the site of the oxideinclusion. FIG. 21 schematically illustrates the coated substrate shownin FIG. 19 wherein the crack formed at the site of the oxide inclusionpropagates through the diffusion boundary and into the substrate. Thesefigures schematically illustrate the propagation of a crack caused by anoxide inclusion in a diffusion coating formed on a substrate. Byremoving the oxide forming materials from the coating composition, suchcracks are reduced or eliminated. Further, the oxide-free coating mayact as a prophylactic preventing the formation of some cracks within thesubstrate.

FIG. 22(a) is a schematic perspective view of a cast turbine engineairfoil part showing a casting defect. FIG. 22(b) is a schematicperspective view of the cast turbine engine airfoil part having the areaof the casting defect being machined. FIG. 22(c) is a schematicperspective view of the cast turbine engine airfoil part after the areaof the casting defect has been machined. FIG. 23(d) is a schematicperspective view of the cast turbine engine airfoil part having the areaof the casting defect being filled with a coating material. FIG. 22(e)is a schematic perspective view of the coated cast turbine engineairfoil part being subjected to a hot isostatic pressing treatment. FIG.22(f) is a schematic perspective view of the repaired cast turbineengine airfoil part.

As shown in FIGS. 23(a) through 23(f), in accordance with this aspect ofthe invention, a method is provided for repairing a turbine engineairfoil part. The dimensional differences are determined betweenpre-repair airfoil dimensions of a turbine engine airfoil part substrateand desired post repair airfoil dimensions of the turbine engine airfoilpart substrate. The pre-repair airfoil dimensions having differentairfoil characteristics than the post-repair airfoil dimensions. Theturbine engine airfoil part being comprised of a metal alloy. The engineairfoil part is coated with a coating capable of forming a diffusionboundary with the turbine engine airfoil part substrate. If necessary,the engine airfoil part can be masked so that only the desired area(s)is coated. The coating material comprises an alloy with substantially nooxide forming constituents so as to avoid the formation of oxideinclusions in the coating material. In addition, or alternatively, themethod of coating can be chosen so as to limit or avoid the formation ofoxide inclusions. For example, the coating can be performed with theairfoil part shrouded with an inert atmosphere, such as argon gas. Or,the coating can be performed under vacuum. In any case, in accordancewith this aspect of the invention, the coating material applied torepair the turbine engine airfoil part is substantially free from oxideinclusions. A hot isostatic heat treatment process is performed toobtain a post-repair turbine engine airfoil part having the desiredpost-repair dimensions and having a substantially oxide free coating anddiffusion bonding between the coating material and the turbine engineairfoil part substrate. The substantially oxide-free coating provides aprotective coating to protect the article from damage. A sintering heattreatment can be performed before the step of performing the hotisostatic heat treatment to limit the occurrence of bubbles on thesurface of the coating material after an isostatic heat treatment. Thesintering heat treatment may be performed at a temperature substantiallythe same as the temperature of the hot isostatic heat treatment.

An aspect of the present invention pertains to a method for forming awear-resistant hard-face contact area on a workpiece, such as a gasturbine engine part. The wear resistant hardface material can be acobalt-based hard facing alloy. The cobalt-based hard facing alloy isparticularly useful as a hard facing material for gas turbine enginecomponents, such as the shrouds of a gas turbine engine blade. Hardfacing material is typically used on the critical components in gasturbine engines that are used to power jet aircraft or for thegeneration of electricity.

FIG. 23(a) is a cut-away view of a gas turbine engine blade showing theshroud portion afixed to the airfoil portion of the blade, and showingthe location of an applied wear resistant hard facing material to thecontact surface of the blade. FIG. 23(b) shows two adjacent blades of anassembled disc showing the contact between the shrouds of two adjacentblades 38 of an assembled disc. The blades are held in the housingmember (not shown) such that surfaces 44 of each shroud section 40contacts corresponding surfaces 44 of adjacent shrouds. These contactsurfaces 44 are subjected to wearing forces during the operation of thegas turbine engine. As an assembled disc of blades rotates, theindividual adjacent blades 38 may chatter against each other, causingwear to occur at the contact surfaces 44 of the shroud sections 40. Thischattering results in constant hammering at the contact surfaces 44 ofthe interlocking blades 38. Excessive wear in the area of the contactsurfaces 44 can have detrimental consequences on the operation of thegas turbine engine. The present invention provides a method for forminga particularly durable and effective hard facing surface to combat theexcessive wear in the area of the contact surfaces of the shrouds. FIG.23(a) shows a typical location for the application of a hard facingmaterial 46. Typically, such hard facing material is applied to theshroud by, for example, manual tig welding or laser welding. Thesemethods result in a mechanical bond being formed between the hard facematerial and the shroud substrate. This mechanical bond is subject tofailure due to chipping or flaking, causing chattering between theshrouds of the assembled disk, and ultimatlely can cause failure of theentire gas turbine engine.

In accordance with the present invention, the wear resistant hardfacesurface is permanently adhered to the shroud substrate through adiffusion bond. Due to the nature of the diffusion bond, as describedherein, the wear resistant hardface material will not chip or flake off,resulting in better service life and possibly will prevent the untimelyfailure of the gas turbine engine and serious detrimental consequences.

In accordance with the invention described in applicants co-pending U.S.patent application, Ser. No. 10/021,107, which is incorporated byreference herein, a cobalt-based alloy is provided that is particularlyuseful as a hard facing material for gas turbine engine components, suchas the shrouds of a gas turbine engine blade. The alloy compositions asdescribed in this co-pending application have a relatively smalllanthanum addition and relatively large carbon content and provideremarkable oxidation resistance and wear resistance at hightemperatures. Importantly, the inventive alloy composition has asuitable combination of ductility and wear resistance at hightemperatures to be effective as a hard face material for limiting theeffects of chattering of blades during the operation of a gas turbineengine. Accordingly, the inventive alloy has a suitable combination ofductility, oxidation resistance and wear resistance and thus representsan improved hard facing material for the blade components of gas turbineengine. The hardface coating material may comprise, for example, analloy characterized by improved oxidation and wear resistance atelevated temperatures consisting essentially in weight percent of about:Percent Carbon 0.07-1.00 Manganese 1.00 Silicon 1.00 Chromium26.00-30.00 Nickel 4.00-6.00 Tungsten 18.00-21.00 Boron  .005-0.100Vanadium 0.75-1.25 Iron 3.00 Lanthanum 0.02-0.12 Cobalt remainder

Alternatively, the hardface coating material may comprise, for example,an alloy characterized by improved oxidation and wear resistance atelevated temperatures consisting essentially in weight percent of about:Percent Carbon 0.08 max Silicon 3.00-3.80 Phosphorus 0.03 max Sulfur0.03 max Chromium 16.50-18.50 Molybdenum 27.00-30.00 Nickel + Iron 3.00max Nitrogen 0.07 max Oxygen 0.05 max Lanthanum 0.02-0.12 Cobaltremainder

FIG. 23(c) is a flowchart illustrating the steps of the inventive methodfor forming a wear-resistant hard-face contact area on a workpiece, suchas a gas turbine engine part. In accordance with the present invention,a method is provided for forming a wear-resistant hardfaced contact areaon the shroud section of a gas turbine engine blade. A predeterminedcontact area of a shroud section of a gas turbine engine blade isselectively coated with a high-density hardface coating material (steptwo). The hardface coating material is capable of forming a diffusionboundary between the hardface coating material and the shroud section. Ahot isostatic heat treatment process is performed to form the diffusionboundary between the hardface coating material and the shroud section toform a wear-resistant hardfaced contact area diffusion bonded to theshroud section (step three).

Depending on the coating process, and the necessity for doing so, thepredetermined contact area can be masked off before the step ofselectively coating (step one). A sintering heat treatment can beperfomed before the step of performing the hot isostatic heat treatmentto limit the occurrence bubbles on the surface of the hardface coatingmaterial after the isostatic heat treatment step (step four). Thesintering heat treatment may be performed at a temperature substantiallythe same as the temperature of the hot isostatic heat treatment. Thehardface coating material may comprise an alloy with substantially nooxide forming constituents so as to avoid the formation of oxideinclusions in the coating material.

With respect to the above description, it is realized that the optimumdimensional relationships for parts of the invention, includingvariations in size, materials, shape, form, function, and manner ofoperation, assembly and use, are deemed readily apparent and obvious toone skilled in the art. All equivalent relationships to thoseillustrated in the drawings and described in the specification areintended to be encompassed by the present invention. Therefore, theforegoing is considered as illustrative only of the principles of theinvention. Further, since numerous modifications and changes willreadily occur to those skilled in the art, it is not desired to limitthe invention to the exact construction and operation shown anddescribed. Accordingly, all suitable modifications and equivalents maybe resorted to, falling within the scope of the invention.

1). A method of forming a wear-resistant hardfaced contact area,comprising the steps of: Selectively coating a predetermined contactarea of a workpiece with a hardface coating material capable of forminga diffusion boundary between the hardface coating material and theworkpiece; and Performing a hot isostatic heat treatment process to formthe diffusion boundary between the hardface coating material and theworkpiece to form a wear-resistant hardfaced contact area diffusionbonded to the workpiece. 2). A method of forming a wear-resistanthardfaced contact area according to claim 1; further comprising the stepof masking off the predetermined contact area before the step ofselectively coating. 3). A method of forming a wear-resistant hardfacedcontact area according to claim 1; further comprising the step ofperforming a sintering heat treatment before the step of performing thehot isostatic heat treatment to limit the occurrence of bubbles on thesurface of the hardface coating material after the isostatic heattreatment step. 4). A method of forming a wear-resistant hardfacedcontact area according to claim 3; wherein the sintering heat treatmentis performed at a temperature substantially the same as the temperatureof the hot isostatic heat treatment. 5). A method of forming awear-resistant hardfaced contact area according to claim 1; wherein thehardface coating material comprises an alloy with substantially no oxideforming constituents so as to avoid the formation of oxide inclusions inthe coating material. 6). A method of forming a wear-resistant hardfacedcontact area according to claim 1; wherein the hardface coating materialcomprise an alloy characterized by improved oxidation and wearresistance at elevated temperatures consisting essentially in weightpercent of about: Percent Carbon 0.07-1.00 Manganese 1.00 Silicon 1.00Chromium 26.00-30.00 Nickel 4.00-6.00 Tungsten 18.00-21.00 Boron .005-0.100 Vanadium 0.75-1.25 Iron 3.00 Lanthanum 0.02-0.12 Cobaltremainder

7). A method of forming a wear-resistant hardfaced contact areaaccording to claim 1; wherein the hardface coating material comprise analloy characterized by improved oxidation and wear resistance atelevated temperatures consisting essentially in weight percent of about:Percent Carbon 0.08 max Silicon 3.00-3.80 Phosphorus 0.03 max Sulfur0.03 max Chromium 16.50-18.50 Molybdenum 27.00-30.00 Nickel + Iron 3.00max Nitrogen 0.07 max Oxygen 0.05 max Lanthanum 0.02-0.12 Cobaltremainder

8). A method of forming a wear-resistant hardfaced contact area on theshroud section of a gas turbine engine blade, comprising the steps of:Selectively coating a predetermined contact area of a shroud section ofa gas turbine engine blade with a hardface coating material capable offorming a diffusion boundary between the hardface coating material andthe shroud section; and Performing a hot isostatic heat treatmentprocess to form the diffusion boundary between the hardface coatingmaterial and the shroud section to form a wear-resistant hardfacedcontact area diffusion bonded to the shroud section. 9). A method offorming a wear-resistant hardfaced contact area on the shroud section ofa gas turbine engine blade according to claim 8; further comprising thestep of masking off the predetermined contact area before the step ofselectively coating. 10). method of forming a wear-resistant hardfacedcontact area on the shroud section of a gas turbine engine bladeaccording to claim 8; further comprising the step of performing asintering heat treatment before the step of performing the hot isostaticheat treatment to limit the occurrence of bubbles on the surface of thehardface coating material after the isostatic heat treatment step. 11).A method of forming a wear-resistant hardfaced contact area on theshroud section of a gas turbine engine blade according to claim 10;wherein the sintering heat treatment is performed at a temperaturesubstantially the same as the temperature of the hot isostatic heattreatment. 12). A method of forming a wear-resistant hardfaced contactarea on the shroud section of a gas turbine engine blade according toclaim 8; wherein the hardface coating material comprises an alloy withsubstantially no oxide forming constituents so as to avoid the formationof oxide inclusions in the coating material. 13). A method of forming awear-resistant hardfaced contact area on the shroud section of a gasturbine engine blade according to claim 8; wherein the hardface coatingmaterial comprise an alloy characterized by improved oxidation and wearresistance at elevated temperatures consisting essentially in weightpercent of about: Percent Carbon 0.07-1.00 Manganese 1.00 Silicon 1.00Chromium 26.00-30.00 Nickel 4.00-6.00 Tungsten 18.00-21.00 Boron .005-0.100 Vanadium 0.75-1.25 Iron 3.00 Lanthanum 0.02-0.12 Cobaltremainder

14). A method of forming a wear-resistant hardfaced contact area on theshroud section of a gas turbine engine blade according to claim 8;wherein the hardface coating material comprise an alloy characterized byimproved oxidation and wear resistance at elevated temperaturesconsisting essentially in weight percent of about: Percent Carbon 0.08max Silicon 3.00-3.80 Phosphorus 0.03 max Sulfur 0.03 max Chromium16.50-18.50 Molybdenum 27.00-30.00 Nickel + Iron 3.00 max Nitrogen 0.07max Oxygen 0.05 max Lanthanum 0.02-0.12 Cobalt remainder